Entry, Descent, and Landing: 2000-2004

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Entry, Descent, and Landing: 2000-2004

This custom bibliography from the NASA Scientific and Technical Information Program lists a sampling of records
found in the NASA Aeronautics and Space Database. The scope of this topic includes technologies for precision
targeting and landing on "high-g" and "low-g" planetary bodies. This area of focus is one of the enabling
technologies as defined by NASA's Report of the President's Commission on Implementation of United States
Space Exploration Policy, published in June 2004.

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Entry, Descent, and Landing
2000-2004

A Custom Bibliography From the

NASA Scientific and Technical Information Program

October 2004

Entry, Descent, and Landing: 2000-2004

This custom bibliography from the NASA Scientific and Technical Information Program lists a sampling of records
found in the NASA Aeronautics and Space Database. The scope of this topic includes technologies for precision
targeting and landing on "high-g" and "low-g" planetary bodies. This area of focus is one of the enabling
technologies as defined by NASA's Report of the President's Commission on Implementation of United States
Space Exploration Policy, published in June 2004.

OCTOBER 2004

20040120953 Computer Sciences Corp., Huntsville, AL, USA
Atmospheric Models for Aerocapture Systems Studies

Justus, C. G.; Duvall, Aleta; Keller, Vernon W.; December 19, 2003; In English, 16-19 Aug. 2004, Providence, RI, USA
Contract(s)/Grant(s): NAS8-60000; No Copyright; Avail: Other Sources

Aerocapture uses atmospheric drag to decelerate into captured orbit from interplanetary transfer orbit. This includes
capture into Earth orbit from, for example, Lunar-return or Mars-return orbit. Eight Solar System destinations have sufficient
atmosphere for aerocapture to be applicable - three of the rocky planets (Venus, Earth, and Mars), four gas giants (Jupiter,
Saturn, Uranus, and Neptune), and Saturn's moon Titan. These destinations fall into two groups: (1) The rocky planets, which
have warm surface temperatures (approx. 200 to 750 K) and rapid decrease of density with altitude, and (2) the gas giants and
Titan, which have cold temperatures (approx. 70 to 170 K) at the surface or 1-bar pressure level, and slow rate of decrease
of density with altitude. The height variation of average density with altitude above 1-bar pressure level for the gas giant
planets is shown. The periapsis density required for aerocapture of spacecraft having typical values of ballistic coefficient (a
measure of mass per unit cross-sectional area) is also shown. The aerocapture altitudes at the gas giants would typically range
from approx. 150 to 300 km. Density profiles are compared for the rocky planets with those for Titan and Neptune.
Aerocapture at the rocky planets would occur at heights of approx. 50 to 100 km. For comparison, typical density and altitudes
for aerobraking operations (circularizing a highly elliptical capture orbit, using multiple atmospheric passes) are also indicated.
Author (revised)
Atmospheric Models; Aerocapture; Aerobraking; Planetary Atmospheres

20040120944 NASA Langley Research Center, Hampton, VA, USA

Preliminary Convective-Radiative Heating Environments for a Neptune Aerocapture Mission

Hollis, Brian R.; Wright, Michael J.; Olejniczak, Joseph; Takashima, Naruhisa; Sutton, Kenneth; Prabhu, Dinesh; [2004]; In
English; AI A A Atmospheric Flight Mechanics Conference and Exhibit, 16-19 Aug. 2004, Providence, RI, USA
Contract(s)/Grant(s): NAS2-99092; NAS1-00135; NCC1-02043; 320-10-00
Report No.(s): AIAA Paper 2004-5177; No Copyright; Avail: CASI; A03, Hardcopy

Convective and radiative heating environments have been computed for a three-dimensional ellipsled configuration which
would perform an aerocapture maneuver at Neptune. This work was performed as part of a one-year Neptune aerocapture
spacecraft systems study that also included analyses of trajectories, atmospheric modeling, aerodynamics, structural design,
and other disciplines. Complementary heating analyses were conducted by separate teams using independent sets of
aerothermodynamic modeling tools (i.e. Navier-Stokes and radiation transport codes). Environments were generated for a
large 5.50 m length ellipsled and a small 2.88 m length ellipsled. Radiative heating was found to contribute up to 80% of the
total heating rate at the ellipsled nose depending on the trajectory point. Good agreement between convective heating
predictions from the two Navier-Stokes solvers was obtained. However, the radiation analysis revealed several uncertainties
in the computational models employed in both sets of codes, as well as large differences between the predicted radiative
heating rates.
Author
Aerocapture; Convective Heat Transfer; Radiative Heat Transfer; Neptune Atmosphere; Aerothermodynamics

20040120869 Morgan Research Corp., Huntsville, AL, USA, NASA Marshall Space Flight Center, Huntsville, AL, USA
Atmospheric Models for Aeroentry and Aeroassist

Justus, C. G; Duvall, Aleta; Keller, Vernon W.; June 15, 2004; In English, 23-26 Aug. 2004, Moffett Field, CA, USA; Original

contains color illustrations

Contract(s)/Grant(s): NNM04AA02C; No Copyright; Avail: CASI; A02, Hardcopy

1

Eight destinations in the Solar System have sufficient atmosphere for aeroentry, aeroassist, or aerobraking/aerocapture:
Venus, Earth, Mars, Jupiter, Saturn, Uranus, and Neptune, plus Saturn's moon Titan. Engineering-level atmospheric models
for Earth, Mars, Titan, and Neptune have been developed for use in NASA s systems analysis studies of aerocapture
applications. Development has begun on a similar atmospheric model for Venus. An important capability of these models is
simulation of quasi-random perturbations for Monte Carlo analyses in developing guidance, navigation and control algorithms,
and for thermal systems design. Characteristics of these atmospheric models are compared, and example applications for
aerocapture are presented. Recent Titan atmospheric model updates are discussed, in anticipation of applications for trajectory
and atmospheric reconstruct of Huygens Probe entry at Titan. Recent and planned updates to the Mars atmospheric model, in
support of future Mars aerocapture systems analysis studies, are also presented.
Author
Atmospheric Models; Solar System; Aeroassist; Aerobraking; Aerocapture; Guidance (Motion)

20040111219 NASA Langley Research Center, Hampton, VA, USA
Structural Design for a Neptune Aerocapture Mission

Dyke, R. Eric; Hrinda, Glenn A.; [2004]; In English, 16-19 Aug. 2004, Providence, RI, USA

Contract(s)/Grant(s): 23-800-90-40

Report No.(s): AIAA Paper 2004-5179; No Copyright; Avail: CASI; A03, Hardcopy

A multi-center study was conducted in 2003 to assess the feasibility of and technology requirements for using aerocapture
to insert a scientific platform into orbit around Neptune. The aerocapture technique offers a potential method of greatly
reducing orbiter mass and thus total spacecraft launch mass by minimizing the required propulsion system mass. This study
involved the collaborative efforts of personnel from Langley Research Center (LaRC), Johnson Space Flight Center (JSFC),
Marshall Space Flight Center (MSFC), Ames Research Center (ARC), and the Jet Propulsion Laboratory (JPL). One aspect
of this effort was the structural design of the full spacecraft configuration, including the ellipsled aerocapture orbiter and the
in-space solar electric propulsion (SEP) module/cruise stage. This paper will discuss the functional and structural requirements
for each of these components, some of the design trades leading to the final configuration, the loading environments, and the
analysis methods used to ensure structural integrity. It will also highlight the design and structural challenges faced while
trying to integrate all the mission requirements. Component sizes, materials, construction methods and analytical results,
including masses and natural frequencies, will be presented, showing the feasibility of the resulting design for use in a Neptune
aerocapture mission. Lastly, results of a post-study structural mass optimization effort on the ellipsled will be discussed,
showing potential mass savings and their influence on structural strength and stiffness
Author
Aerocapture; Neptune (Planet); Structural Design; Space Missions; Spacecraft Design

20040111218 NASA Langley Research Center, Hampton, VA, USA
Aerocapture Performance Analysis for a Neptune-Triton Exploration Mission

Starr, Brett R.; Westhelle, Carlos H.; Masciarelli, James P.; [2004]; In English, 16-19 Aug. 2004, Providence, RI, USA

Contract(s)/Grant(s): 23-800-90-50

Report No.(s): AIAA Paper 2004-4955; No Copyright; Avail: CASI; A03, Hardcopy

A systems analysis has been conducted for a Neptune-Triton Exploration Mission in which aerocapture is used to capture
a spacecraft at Neptune. Aerocapture uses aerodynamic drag instead of propulsion to decelerate from the interplanetary
approach trajectory to a captured orbit during a single pass through the atmosphere. After capture, propulsion is used to move
the spacecraft from the initial captured orbit to the desired science orbit. A preliminary assessment identified that a spacecraft
with a lift to drag ratio of 0.8 was required for aerocapture. Performance analyses of the 0.8 L/D vehicle were performed using
a high fidelity flight simulation within a Monte Carlo executive to determine mission success statistics. The simulation was
the Program to Optimize Simulated Trajectories (POST) modified to include Neptune specific atmospheric and planet models,
spacecraft aerodynamic characteristics, and interplanetary trajectory models. To these were added autonomous guidance and
pseudo flight controller models. The Monte Carlo analyses incorporated approach trajectory delivery errors, aerodynamic
characteristics uncertainties, and atmospheric density variations. Monte Carlo analyses were performed for a reference set of
uncertainties and sets of uncertainties modified to produce increased and reduced atmospheric variability. For the reference
uncertainties, the 0.8 L/D flatbottom ellipsled vehicle achieves 100% successful capture and has a 99.87 probability of
attaining the science orbit with a 360 m/s V budget for apoapsis and periapsis adjustment. Monte Carlo analyses were also
performed for a guidance system that modulates both bank angle and angle of attack with the reference set of uncertainties.

An alpha and bank modulation guidance system reduces the 99.87 percentile DELTA V 173 m/s (48%) to 187 m/s for the

reference set of uncertainties.

Author

Aerocapture; Neptune (Planet); Space Missions; Triton; Systems Analysis; Spacecraft Performance; Space Exploration

20040111217 NASA Langley Research Center, Hampton, VA, USA
Neptune Aerocapture Systems Analysis

Lockwood, Mary Kae; [2004]; In English, 16-19 Aug. 2004, Providence, RI, USA

Contract(s)/Grant(s): 23-800-90-10

Report No.(s): AIAA Paper 2004-4951; No Copyright; Avail: CASI; A03, Hardcopy

A Neptune Aerocapture Systems Analysis is completed to determine the feasibility, benefit and risk of an aeroshell
aerocapture system for Neptune and to identify technology gaps and technology performance goals. The high fidelity systems
analysis is completed by a five center NASA team and includes the following disciplines and analyses: science; mission
design; aeroshell configuration screening and definition; interplanetary navigation analyses; atmosphere modeling;
computational fluid dynamics for aerodynamic performance and database definition; initial stability analyses; guidance
development; atmospheric flight simulation; computational fluid dynamics and radiation analyses for aeroheating environment
definition; thermal protection system design, concepts and sizing; mass properties; structures; spacecraft design and
packaging; and mass sensitivities. Results show that aerocapture can deliver 1.4 times more mass to Neptune orbit than an
all-propulsive system for the same launch vehicle. In addition aerocapture results in a 3-4 year reduction in trip time compared
to all-propulsive systems. Aerocapture is feasible and performance is adequate for the Neptune aerocapture mission. Monte
Carlo simulation results show 100% successful capture for all cases including conservative assumptions on atmosphere and
navigation. Enabling technologies for this mission include TPS manufacturing; and aerothermodynamic methods and
validation for determining coupled 3-D convection, radiation and ablation aeroheating rates and loads, and the effects on
surface recession.
Author
Aerocapture; Neptune (Planet); Systems Analysis; Technology Utilization; Aeroshells

20040095913 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA, USA
Mars Exploration Rovers Landing Dispersion Analysis

Knocke, Philip C; Wawrzyniak, Geoffrey G.; Kennedy, Brian M.; Desai, Prasun N.; Parker, Timothy J.; Golombek, Matthew

P.; Duxbury, Thomas C; Kass, David M.; [2004]; In English

Report No.(s): AIAA Paper 2004-5093; No Copyright; Avail: CASI; A03, Hardcopy

Landing dispersion estimates for the Mars Exploration Rover missions were key elements in the site targeting process and
in the evaluation of landing risk. This paper addresses the process and results of the landing dispersion analyses performed
for both Spirit and Opportunity. The several contributors to landing dispersions (navigation and atmospheric uncertainties,
spacecraft modeling, winds, and margins) are discussed, as are the analysis tools used. JPL's MarsLS program, a
MATLAB-based landing dispersion visualization and statistical analysis tool, was used to calculate the probability of landing
within hazardous areas. By convolving this with the probability of landing within flight system limits (in-spec landing) for
each hazard area, a single overall measure of landing risk was calculated for each landing ellipse. In-spec probability contours
were also generated, allowing a more synoptic view of site risks, illustrating the sensitivity to changes in landing location, and
quantifying the possible consequences of anomalies such as incomplete maneuvers. Data and products required to support
these analyses are described, including the landing footprints calculated by NASA Langley' s POST program and JPL's AEPL
program, cartographically registered base maps and hazard maps, and flight system estimates of in-spec landing probabilities
for each hazard terrain type. Various factors encountered during operations, including evolving navigation estimates and
changing atmospheric models, are discussed and final landing points are compared with approach estimates.
Author
Mars Exploration; Landing Modules; Roving Vehicles; Statistical Analysis; Land

20040095912 NASA Langley Research Center, Hampton, VA, USA, Jet Propulsion Lab., California Inst, of Tech., Pasadena,

CA, USA

Mars Exploration Rovers Entry, Descent, and Landing Trajectory Analysis

Desai, Prasun N.; Knocke, Philip C; August 11, 2004; In English, 16-19 Aug. 2004, Providence, RI, USA

Contract(s)/Grant(s): 23-749-30-00

Report No.(s): AIAA Paper 2004-5092; No Copyright; Avail: CASI; A02, Hardcopy

The Mars Exploration Rover mission successfully landed two rovers 'Spirit' and 'Opportunity' on Mars on January 4th
and 25th of 2004, respectively. The trajectory analysis performed to define the entry, descent, and landing (EDL) scenario is
described. The entry requirements and constraints are presented, as well as uncertainties used in a Monte Carlo dispersion
analysis to statistically assess the robustness of the entry design to off-nominal conditions. In the analysis, six-degree-of-
freedom and three-degree-of-freedom trajectory results are compared to assess the entry characteristics of the capsule.
Comparison of the preentry results to preliminary post-landing reconstruction data shows that all EDL parameters were within
the requirements. In addition, the final landing position for both 'Spirit' and 'Opportunity' were within 15 km of the predicted
landing location.
Author
Trajectory Analysis; Mars Landing; Roving Vehicles; Position (Location)

20040086474 NASA Langley Research Center, Hampton, VA, USA
Aeroassist Technology Planning for Exploration

Munk, Michelle M.; Powell, Richard W.; [2000]; In English

Report No.(s): AAS-00-169; No Copyright; Avail: CASI; A03, Hardcopy

Now that the International Space Station is undergoing assembly, NASA is strategizing about the next logical exploration
strategy for robotic missions and the next destination for humans. NASA's current efforts are in developing technologies that
will both aid the robotic exploration strategy and make human flight to other celestial bodies both safe and affordable. One
of these enabling technologies for future robotic and human exploration missions is aeroassist. This paper will (1) define
aeroassist, (2) explain the benefits and uses of aeroassist, and (3) describe a method, currently used by the NASA Aeroassist
Working Group, by which widely geographically distributed teams can assemble, present, use, and archive technology
information.
Author
Aeroassist; International Space Station; NASA Space Programs; Space Exploration; Technological Forecasting

20040085958 Morgan Research Corp., Huntsville, AL, USA, NASA Marshall Space Flight Center, Huntsville, AL, USA
Atmospheric Models for Aerocapture

Justus, C. G.; Duvall, Aleta L.; Keller, Vernon W.; April 09, 2004; In English, 11-14 Jul. 2004, Fort Lauderdale, FL, USA
Contract(s)/Grant(s): NNM04AA02C; No Copyright; Avail: CASI; A02, Hardcopy

There are eight destinations in the solar System with sufficient atmosphere for aerocapture to be a viable aeroassist option
- Venus, Earth, Mars, Jupiter, Saturn and its moon Titan, Uranus, and Neptune. Engineering-level atmospheric models for four
of these targets (Earth, Mars, Titan, and Neptune) have been developed for NASA to support systems analysis studies of
potential future aerocapture missions. Development of a similar atmospheric model for Venus has recently commenced. An
important capability of all of these models is their ability to simulate quasi-random density perturbations for Monte Carlo
analyses in developing guidance, navigation and control algorithm, and for thermal systems design. Similarities and
differences among these atmospheric models are presented, with emphasis on the recently developed Neptune model and on
planned characteristics of the Venus model. Example applications for aerocapture are also presented and illustrated. Recent
updates to the Titan atmospheric model are discussed, in anticipation of applications for trajectory and atmospheric reconstruct
of Huygens Probe entry at Titan.
Author
Atmospheric Models; Aerocapture; Huygens Probe; Environmental Monitoring

20040085708 NASA Langley Research Center, Hampton, VA, USA
Aeroheating Thermal Analysis Methods for Aerobraking Mars Missions

Amundsen, Ruth M.; Dec, John A.; George, Benjamin E.; [2002]; In English; No Copyright; Avail: CASI; A03, Hardcopy
Mars missions often employ aerobraking upon arrival at Mars as a low-mass method to gradually reduce the orbit period
from a high-altitude, highly elliptical insertion orbit to the final science orbit. Two recent missions that made use of
aerobraking were Mars Global Surveyor (MGS) and Mars Odyssey. Both spacecraft had solar arrays as the main aerobraking
surface area. Aerobraking produces a high heat load on the solar arrays, which have a large surface area exposed to the airflow
and relatively low mass. To accurately model the complex behavior during aerobraking, the thermal analysis must be tightly
coupled to the flight mechanics, aerodynamics, and atmospheric modeling efforts being performed during operations. To
properly represent the temperatures prior to and during the drag pass, the model must include the orbital solar and planetary
heat fluxes. The correlation of the thermal model to flight data allows a validation of the modeling process, as well as

information on what processes dominate the thermal behavior. This paper describes the thermal modeling method that was

developed for this purpose, as well as correlation for two flight missions, and a discussion of improvements to the

methodology.

Author

Aerobraking; Mars Missions; Elliptical Orbits; Thermal Analysis; Aerodynamic Heating

20040068067 Computer Sciences Corp., Huntsville, AL, USA

Connecting Atmospheric Science and Atmospheric Models for Aerocaptured Missions to Titan and the Outer Planets

Justus, C. G.; Duvall, Aleta; Keller, Vernon W.; December 19, 2003; In English, 25-30 Apr. 2004, Nice, France
Contract(s)/Grant(s): NAS8-60000; No Copyright; Avail: Other Sources; Abstract Only

Many atmospheric measurement systems, such as the sounding instruments on Voyager, gather atmospheric information
in the form of temperature versus pressure level. In these terms, there is considerable consistency among the mean atmospheric
profiles of the outer planets Jupiter through Neptune, including Titan. On a given planet or on Titan, the range of variability
of temperature versus pressure level due to seasonal, latitudinal, and diurnal variations is also not large. However, many
engineering needs for atmospheric models relate not to temperature versus pressure level but atmospheric density versus
geometric altitude. This need is especially true for design and analysis of aerocapture systems. Aerocapture drag force
available for aerocapture is directly proportional to atmospheric density. Available aerocapture 'corridor width' (allowable
range of atmospheric entry angle) also depends on height rate of change of atmospheric density, as characterized by density
scale height. Characteristics of hydrostatics and the gas law equation mean that relatively small systematic differences in
temperature-versus-pressure profiles can integrate at high altitudes to very large differences in density-versus-altitude profiles.
Thus a given periapsis density required to accomplish successful aerocapture can occur at substantially different altitudes
(approx. 150 - 300 km) on the various outer planets, and significantly different density scale heights (approx. 20 - 50 km) can
occur at these periapsis altitudes. This paper will illustrate these effects and discuss implications for improvements in
atmospheric measurements to yield significant impact on design of aerocapture systems for future missions to Titan and the
outer planets. Relatively small- scale atmospheric perturbations, such as gravity waves, tides, and other atmospheric variations
can also have significant effect on design details for aerocapture guidance and control systems. This paper will also discuss
benefits that would result from improved understanding of Titan and outer planetary atmospheric perturbation characteristics.
Details of recent engineering-level atmospheric models for Titan and Neptune will be presented, and effects of present and
future levels of atmospheric uncertainty and variability characteristics will be examined.
Author
Atmospheric Physics; Atmospheric Models; Aerocapture; Planetary Atmospheres; Atmospheric Density; Annual Variations

20040062499 Lunar and Planetary Inst., Houston, TX, USA

Lunar and Planetary Science XXXV: Missions and Instruments: Hopes and Hope Fulfilled

2004; In English; Lunar and Planetary Science XXXV, 15-19 Mar. 2004, Houston, TX, USA

Contract(s)/Grant(s): NCC5-679

Report No.(s): LPI-Contrib-1197; Copyright; Avail: CASI; C01, CD-ROM

The titles in this section include: 1) Mars Global Surveyor Mars Orbiter Camera in the Extended Mission: The MOC
Toolkit; 2) Mars Odyssey THEMIS-VIS Calibration; 3) Early Science Operations and Results from the ESA Mars Express
Mission: Focus on Imaging and Spectral Mapping; 4) The Mars Express/NASA Project at JPL; 5) Beagle 2: Mission to Mars
- Current Status; 6) The Beagle 2 Microscope; 7) Mars Environmental Chamber for Dynamic Dust Deposition and Statics
Analysis; 8) Locating Targets for CRISM Based on Surface Morphology and Interpretation of THEMIS Data; 9) The Phoenix
Mission to Mars; 10) First Studies of Possible Landing Sites for the Phoenix Mars Scout Mission Using the BMST; 11) The
2009 Mars Telecommunications Orbiter; 12) The Aurora Exploration Program - The ExoMars Mission; 13) Electron-induced
Luminescence and X-Ray Spectrometer (ELXS) System Development; 14) Remote-Raman and Micro-Raman Studies of Solid
C02, CH4, Gas Hydrates and Ice; 15) The Compact Microimaging Spectrometer (CMIS): A New Tool for In-Situ Planetary
Science; 16) Preliminary Results of a New Type of Surface Property Measurement Ideal for a Future Mars Rover Mission;
17) Electrodynamic Dust Shield for Solar Panels on Mars; 18) Sensor Web for Spatio-Temporal Monitoring of a Hydrological
Environment; 19) Field Testing of an In-Situ Neutron Spectrometer for Planetary Exploration: First Results; 20) A Miniature
Solid-State Spectrometer for Space Applications - Field Tests; 21) Application of Laser Induced Breakdown Spectroscopy
(LIBS) to Mars Polar Exploration: LIBS Analysis of Water Ice and Water Ice/Soil Mixtures; 22) LIBS Analysis of Geological
Samples at Low Pressures: Application to Mars, the Moon, and Asteroids; 23) In-Situ 1-D and 2-D Mapping of Soil Core and
Rock Samples Using the LIBS Long Spark; 24) Rocks Analysis at Stand Off Distance by LIBS in Martian Conditions; 25)
Evaluation of a Compact Spectrograph/Detection System for a LIBS Instrument for In-Situ and Stand-Off Detection; 26)

Analysis of Organic Compounds in Mars Analog Samples; 27) Report of the Organic Contamination Science Steering Group;
28) The Water- Wheel IR (WIR) - A Contact Survey Experiment for Water and Carbonates on Mars; 29) Mid-IR Fiber Optic
Probe for In Situ Water Detection and Characterization; 30) Effects of Subsurface Sampling & Processing on Martian Simulant
Containing Varying Quantities of Water; 31) The Subsurface Ice Probe (SIPR): A Low-Power Thermal Probe for the Martian
Polar Layered Deposits; 32) Deploying Ground Penetrating Radar in Planetary Analog Sites to Evaluate Potential Instrument
Capabilities on Future Mars Missions; 33) Evaluation of Rock Powdering Methods to Obtain Fine-grained Samples for
CHEMIN, a Combined XRD/XRF Instrument; 34) Novel Sample-handling Approach for XRD Analysis with Minimal Sample
Preparation; 35) A New Celestial Navigation Method for Mars Landers; 36) Mars Mineral Spectroscopy Web Site: A Resource
for Remote Planetary Spectroscopy.
CASI
Spacecraft Instruments; Planetology; Mars Missions

20040039671

Multibody Parachute Flight Simulations for Planetary Entry Trajectories Using 'Equilibrium Points'

Raiszadeh, Ben; Advances in the Astronautical Sciences; 2003; ISSN 0065-3438; Volume 114, Issue SUPPL., p. 903-914; In
English; Copyright; Avail: Other Sources

A method has been developed to reduce numerical stiffness and computer CPU requirements of high fidelity multibody
flight simulations involving parachutes for planetary entry trajectories. Typical parachute entry configurations consist of entry
bodies suspended from a parachute, connected by flexible lines. To accurately calculate line forces and moments, the
simulations need to keep track of the point where the flexible lines meet (confluence point). In previous multibody parachute
flight simulations, the confluence point has been modeled as a point mass. Using a point mass for the confluence point tends
to make the simulation numerically stiff, because its mass is typically much less that than the main rigid body masses. One
solution for stiff differential equations is to use a very small integration time step. However, this results in large computer CPU
requirements. In the method described in the paper, the need for using a mass as the confluence point has been eliminated.
Instead, the confluence point is modeled using an 'equilibrium point' . This point is calculated at every integration step as the
point at which sum of all line forces is zero (static equilibrium). The use of this 'equilibrium point' has the advantage of both
reducing the numerical stiffness of the simulations, and eliminating the dynamical equations associated with vibration of a
lumped mass on a high-tension string.
EI
Computers; Parachutes; Spacecraft; Stiffness; Trajectories

20040039371

Approach navigation for the 2009 Mars large lander

Burkhart, P. Daniel; Advances in the Astronautical Sciences; 2003; ISSN 0065-3438; Volume 114, Issue SUPPL., p.
2181-2196; In English; Copyright; Avail: Other Sources

The current Mars exploration plan envisions the launch of a large lander in the 2009 launch opportunity with a soft landing
on Mars in the fall of 2010. The goal is to achieve a landed surface position within 10km of the target landing site. Current
entry descent and landing (EDL) analysis shows that the largest contributor to the landed position error is uncertainty of the
initial conditions, which are supplied by the ground-based navigation process. The focus of this paper is the performance of
the approach navigation process using combinations of Deep Space Network (DSN) Doppler, ranging and delta differential
one-way range (delta DOR) measurements along with optical navigation data collected by the spacecraft. Results for several
combinations of data types will be included.
EI
Data Acquisition; Navigation; Planetary Landing; Roving Vehicles; Trajectories

20040039331

Optical landmark detection for spacecraft navigation

Cheng, Yang; Johnson, Andrew E.; Matthies, Larry H.; Olson, Clark F; Advances in the Astronautical Sciences; 2003; ISSN
0065-3438; Volume 114, Issue SUPPL., p. 1767-1785; In English; Copyright; Avail: Other Sources

Optical landmark navigation using craters on the surface of a central body was first used operationally by the Near Earth
Asteroid Rendezvous (NEAR) mission. It has proven to be a powerful data type for determining spacecraft orbits above the
target for close flybys and low altitude orbiting. Tracking individual landmarks, which are small craters, enables orbit
determination accuracies on the order of the camera resolution or several meters. This exceeds the accuracy that can be

obtained from radiometric data alone. Currently, most of optical landmark navigation operations, such as crater detection,
tracking, and matching etc, are done manually, which is extremely time consuming, tedious and sometime unmanageable.
Because of the lengthily operation time and the deep-space communication delay, manual operation cannot meet the
requirements of rapid and precise spacecraft maneuvers such as close orbiting, fast fiybys and landing. Automating this
operation can greatly improve navigation accuracy and efficiency and ultimately lead to an on-board autonomous navigation
capability. In this paper, a new crater detection algorithm is suggested. Experimental studies show that this new algorithm can
achieve sub-pixel accuracy in position, its detection rate is better than 90% and its false alarm rate is less than 5%. These good
characteristics indicate that it is an ideal crater detection algorithm for spacecraft optical navigation.
EI
Cameras; Navigation; Resolution; Spacecraft Propulsion; Tracking (Position)

20040039275

Daily repeat-groundtrack Mars orbits

Noreen, Gary; Kerridge, Stuart; Diehl, Roger; Neelon, Joseph; Ely, Todd; Turner, Andrew E.; Advances in the Astronautical
Sciences; 2003; ISSN 0065-3438; Volume 114, Issue SUPPL., p. 1143-1155; In English; Copyright; Avail: Other Sources

This paper derives orbits at Mars with groundtracks that repeat at the same times every solar day (sol). A relay orbiter
in such an orbit would pass over in-situ probes at the same times every sol, ensuring consistent coverage and simplifying
mission design and operations. 42 orbits in five classes are characterized: 14 circular equatorial prograde orbits 14 circular
equatorial retrograde orbits 1 1 circular sun synchronous orbits 2 eccentric equatorial orbits 1 eccentric critically inclined orbit
The paper reports on the performance of a relay orbiter in some of the orbits.
EI
Aerospace Sciences; Communication Satellites; Ground Tracks; Orbits; Planetary Landing; Planets

20040038205

Entry trajectory and atmosphere reconstruction methodologies for the mars exploration rover mission

Desai, Prasun N.; Blanchard, Robert C; Powell, Richard W.; European Space Agency, (Special Publication) ESA SP; February
2004; ISSN 0379-6566, Issue no. 544, p. 213-220; In English; International Workshop: Planetary Probe Atmospheric Entry
and Descent Trajectory Analysis and Science, Oct. 6-9, 2003, Lisbon, Portugal; Copyright; Avail: Other Sources

The Mars Exploration Rover (MER) mission will land two landers on the surface of Mars, arriving in January 2004. Both
landers will deliver the rovers to the surface by decelerating with the aid of an aeroshell, a supersonic parachute, retro-rockets,
and air bags for safely landing on the surface. The reconstruction of the MER descent trajectory and atmosphere profile will
be performed for all the phases from hypersonic flight through landing. A description of multiple methodologies for the flight
reconstruction is presented from simple parameter identification methods through a statistical Kalman filter approach.
EI
Air Bag Restraint Devices; Kalman Filters; Parachutes; Planetary Landing; Trajectories

20040038193

Entry descent, and landing scenario for the Mars exploration Rover mission

Desai, Prasun N; Lee, Wayne J.; European Space Agency, (Special Publication) ESA SP; February 2004; ISSN 0379-6566,
Issue no. 544, p. 31-36; In English; International Workshop: Planetary Probe Atmospheric Entry and Descent Trajectory
Analysis and Science, Oct. 6-9, 2003, Lisbon, Portugal; Copyright; Avail: Other Sources

In January 2004, the Mars Exploration Rover (MER) mission will land two landers on the surface of Mars. Both landers
will deliver a rover to the surface using an entry, descent, and landing (EDL) scenario based on Mars Pathfinder heritage.
However, the entry conditions and environments are different from that of Mars Pathfinder. Unique challenges are present due
to the entry differences of a heavier entry mass, less dense atmosphere, and higher surface landing site altitude. These
differences result in a higher terminal velocity and less time for performing all the EDL events as compared to Mars Pathfinder.
As a result of these differences, modifications are made to the MER EDL systems to safely deliver the rovers to the surface
of Mars.
EI
Aerospace Sciences; Planetary Landing; Planets; Roving Vehicles; Topography

20040038111

Thermal protection system technology and facility needs for demanding future planetary missions

Laub, B.; Venkatapathy, E.; European Space Agency, (Special Publication) ESA SP; February 2004; ISSN 0379-6566, Issue
no. 544, p. 239-247; In English; International Workshop: Planetary Probe Atmospheric Entry and Descent Trajectory Analysis
and Science, Oct. 6-9, 2003, Lisbon, Portugal; Copyright; Avail: Other Sources

NASA has successfully launched numerous science missions to inner and outer planets in our solar system of which the
most challenging were to Venus and Jupiter and the knowledge gained from those missions have been invaluable yet
incomplete. Future missions will be built on what we have learned from the past missions but they will be more demanding
from both the science as well as the mission design and engineering perspectives. The Solar System Exploration Decadal
Survey (SSEDS) produced for NASA by the National Research Council identified a broad range of science objectives many
of which can only be satisfied with atmospheric entry probes. The SSEDS recommended new probe/lander missions to both
Venus and Jupiter. The Pioneer- Venus probe mission was launched in August 1978 and four probes successfully entered the
Venusian atmosphere in December 1978. The Galileo mission was launched in October 1989 and one probe successfully
entered the Jovian atmosphere in December 1995. The thermal protection system requirements for these two missions were
unlike any other planetary probes and required fully dense carbon phenolic for the forebody heat shield. Developing thermal
protection systems to accomplish future missions outlined in the Decadal Survey presents a technology challenge since they
will be more demanding than these past missions. Unlike Galileo, carbon phenolic may not be an adequate TPS for a future
Jupiter multiprobe mission since non-equatorial probes will enter at significantly higher velocity than the Galileo equatorial
probe and the entry heating scales approximately with the cube of the entry velocity. At such heating rates the TPS mass
fraction for a carbon phenolic heat shield would be prohibitive. A new, robust and efficient TPS is required for such probes.
The Giant Planet Facility (GPF), developed and employed during the development of the TPS for the Galileo probe was
dismantled after completion of the program. Furthermore, flight data from the Galileo probe suggested that the complex
physics associated with the interaction between massive ablation and a severe shock layer radiation environment is not well
understood or modeled. The lack of adequate ground test facilities to support the development and qualification of new TPS
materials adds additional complexities. The requirements for materials development, ground testing and sophisticated
modeling to enable these challenging missions are the focus of this paper.
EI
Aerospace Sciences; Heat Shielding; Planetary Landing; Space Probes; Vaporizing

20040038093

Ultra-stable oscillators for planetary entry probes

Asmar, S. W.; Atkinson, D. H.; Bird, M. K.; Wood, G. E.; European Space Agency, (Special Publication) ESA SP; February
2004; ISSN 0379-6566, Issue no. 544, p. 131-134; In English; International Workshop: Planetary Probe Atmospheric Entry
and Descent Trajectory Analysis and Science, Oct. 6-9, 2003, Lisbon, Portugal; Copyright; Avail: Other Sources

Ultra-stable oscillators on-board planetary missions were developed for Radio Science instrumentation, functioning as
frequency references for the one-way downlink during atmospheric occultations. They have also been flown on planetary entry
probes including the Jupiter entry probe, carried by Galileo, and the Huygens Titan entry probe, carried by Cassini, for
performing Doppler Wind Experiments. The Jupiter and Titan probes utilized different oscillators, quartz and rubidium,
respectively. This paper presents the development of ultra-stable oscillators on deep space missions and discusses the tradeoffs
encountered when selecting oscillators for planetary entry probes, including factors such as duration of the experiment, the
available warm-up time and the Allan deviation and phase noise requirements.
EI
Mechanical Oscillators; Quartz; Rubidium; Transponders

20040038075

Pioneer Venus and Galileo entry probe heritage

Bienstock, Bernard J.; European Space Agency, (Special Publication) ESA SP; February 2004; ISSN 0379-6566, Issue no.
544, p. 37-45; In English; International Workshop: Planetary Probe Atmospheric Entry and Descent Trajectory Analysis and
Science, Oct. 6-9, 2003, Lisbon, Portugal; Copyright; Avail: Other Sources

Beginning in the late 1960s, NASA began planning for its first program to explore Venus. Although planetary entry probes
had been flown to Venus by the Soviets beginning in 1967, NASA had not previously flown this type of mission, The Space
and Communications Group of Hughes Aircraft Company, now owned by Boeing and called Boeing Satellite Systems, worked
with NASA to perform initial studies that culminated with a contract for the Pioneer Venus program in early 1974, Pioneer
Venus was an ambitious program that included four planetary entry probes, transported to Venus by a Multiprobe Bus, and

a Venus Orbiter. This paper focuses on the engineering aspects of the probes and the challenges overcome in accommodating
the various scientific instruments. The second NASA planetary entry program was the Galileo Mission that began with initial
studies in the early 1970s. This mission to Jupiter included both an Orbiter and a Probe. Although the Galileo Probe planetary
entry program was begun as the Pioneer Venus probes were heading towards Venus, there were significant engineering
differences between the Pioneer Venus probe designs and the Galileo Probe. These differences, dictated by a number of factors,
are discussed. The paper concludes with a summary of lessons learned by Boeing and NASA in designing, manufacturing and
ultimately flying the Venus and Jupiter planetary entry probes.
EI
Aerospace Sciences; Galileo Spacecraft; Planets; Pressure Vessels; Space Probes; Venus (Planet)

20040038071

Summary of the Boulder Entry Probe Workshop April 21-22, 2003, Boulder, Colorado, USA

Young, Richard E.; Atkinson, David; Atreya, Sushil; Banfield, Donald; Beebe, Reta; Bolton, Scott; Briggs, Geoffrey; Crisp,
David; Cutts, James; Drake, Michael; Esposito, Larry; Galal, Kenneth; Hubbard, William; Hunten, Donald; Ingersoll, Andrew;
et al., T; European Space Agency, (Special Publication) ESA SP; February 2004; ISSN 0379-6566, Issue no. 544, p. 13-20;
In English; International Workshop: Planetary Probe Atmospheric Entry and Descent Trajectory Analysis and Science, Oct.
6-9, 2003, Lisbon, Portugal; Copyright; Avail: Other Sources

The Solar System Exploration Decadal Survey (SSEDS) identified several high priority Solar System Key Science
Questions that should be addressed by entry probes/landers, or that should be addressed simultaneously by both probes/landers
and remote sensing types of observations. These Key Science Questions are directly relevant to Goals and Objectives of the
NASA Strategic Plan and Office of Space Science Strategic Plan. In this report we define entry probes/landers as spacecraft
that sample in-situ a planetary atmosphere, and planetary surface if there is one. The Entry Probe Workshop grew out of a
community concern that if entry probes/landers were to be a viable mission option for addressing the overarching questions
identified in the SSEDS in the coming 10-15 years, significant resources must be applied to key technology areas immediately.
The major science objectives requiring entry probes and the key technology development areas for probes are described.
EI
Aerospace Sciences; Meteorites; Periodic Variations; Planetary Landing; Planets; Space Probes

20040037789 NASA Langley Research Center, Hampton, VA, USA
Angle-of-Attack-Modulated Terminal Point Control for Neptune Aerocapture

Queen, Eric M.; [2004]; In English, 8-12 Feb. 2004, Maui, HI, USA

Contract(s)/Grant(s): 800-90-50

Report No.(s): AAS-04-129; Copyright; Avail: CASI; A02, Hardcopy

An aerocapture guidance algorithm based on a calculus of variations approach is developed, using angle of attack as the
primary control variable. Bank angle is used as a secondary control to alleviate angle of attack extremes and to control
inclination. The guidance equations are derived in detail. The controller has very small onboard computational requirements
and is robust to atmospheric and aerodynamic dispersions. The algorithm is applied to aerocapture at Neptune. Three versions
of the controller are considered with varying angle of attack authority. The three versions of the controller are evaluated using
Monte Carlo simulations with expected dispersions.
Author
Algorithms; Aerocapture; Angle of Attack; Neptune (Planet); Control Theory; Terminal Guidance

20040037788 NASA Langley Research Center, Hampton, VA, USA

Mars Exploration Rover Terminal Descent Mission Modeling and Simulation

Raiszadeh, Behzad; Queen, Eric M.; February 2004; In English, 8-12 Feb. 2004, Maui, HI, USA

Contract(s)/Grant(s): 759-30-00

Report No.(s): AAS-04-271; No Copyright; Avail: CASI; A03, Hardcopy

Because of NASA's added reliance on simulation for successful interplanetary missions, the MER mission has developed
a detailed EDL trajectory modeling and simulation. This paper summarizes how the MER EDL sequence of events are
modeled, verification of the methods used, and the inputs. This simulation is built upon a multibody parachute trajectory
simulation tool that has been developed in POST II that accurately simulates the trajectory of multiple vehicles in flight with
interacting forces. In this model the parachute and the suspended bodies are treated as 6 Degree-of-Freedom (6 DOF) bodies.
The terminal descent phase of the mission consists of several Entry, Descent, Landing (EDL) events, such as parachute

deployment, heatshield separation, deployment of the lander from the backshell, deployment of the airbags, RAD firings, TIRS

firings, etc. For an accurate, reliable simulation these events need to be modeled seamlessly and robustly so that the

simulations will remain numerically stable during Monte-Carlo simulations. This paper also summarizes how the events have

been modeled, the numerical issues, and modeling challenges.

Author

Mars Exploration; Mars Roving Vehicles; Descent; Space Missions; Mathematical Models; Trajectory Analysis;

Computerized Simulation

20040024535

Planning for a Mars in situ sample preparation and distribution (SPAD) system

Beaty, D. W.; Miller, S.; Zimmerman, W.; Bada, J.; Conrad, P.; Dupuis, E.; Huntsberger, T.; Ivlev, R.; Kim, S. S.; Lee, B. G.;
Lindstrom, D.; Lorenzoni, L.; Mahaffy, P.; McNamara, K.; Papanastassiou, D.; et al., T; Planetary and Space Science;
January/March 2004; ISSN 0032-0633; Volume 52, Issue no. 1-3, p. 55-66; In English; Copyright; Avail: Other Sources

For Mars in situ landed missions, it has become increasingly apparent that significant value may be provided by a shared
system that we call a Sample Preparation and Distribution (SPAD) System. A study was conducted to identify the issues and
feasibility of such a system for these missions that would provide common functions for: receiving a variety of sample types
from multiple sample acquisition systems; conducting preliminary characterization of these samples with non-destructive
science instruments and making decisions about what should happen to the samples; performing a variety of sample
preparation functions; and, finally, directing the prepared samples to additional science instruments for further analysis.
Scientific constraints on the functionality of the system were identified, such as triage, contamination management, and
various sample preparation steps, e.g., comminution, splitting, rock surfacing, and sieving. Some simplifying strategies were
recommended and an overall science flow was developed. Engineering functional requirements were also investigated and
example architectures developed. Preliminary conclusions are that shared SPAD facility systems could indeed add value to
future Mars in situ landed missions if they are designed to respond to the particular requirements and constraints of those
missions, that such a system appears feasible for consideration, and that certain standards should be developed for key SPAD
interfaces, (copyright) 2003 Elsevier Ltd. All rights reserved.
EI
Aerospace Sciences; In Situ Measurement; Planetary Landing; Spacecraft

20040024261

Blended control, predictor-corrector guidance algorithm: An enabling technology for Mars aerocapture

Jits, Roman Y.; Walberg, Gerald D.; Acta Astronautica; March 2004; ISSN 0094-5765; Volume 54, Issue no. 6, p. 385-398;
In English; Copyright; Avail: Other Sources

A guidance scheme designed for coping with significant dispersion in the vehicle's state and atmospheric conditions is
presented. In order to expand the flyable aerocapture envelope, control of the vehicle is realized through bank angle and
angle-of-attack modulation. Thus, blended control of the vehicle is achieved, where the lateral and vertical motions of the
vehicle are decoupled. The overall implementation approach is described, together with the guidance algorithm macrologic
and structure. Results of guidance algorithm tests in the presence of various single and multiple off-nominal conditions are
presented and discussed, (copyright) 2003 Published by Elsevier Ltd.
EI
Aerospace Sciences; Astrophysics; Atmospheric Chemistry; Interplanetary Spacecraft; Planets; Predictor-Corrector Methods

20040012726 NASA Marshall Space Flight Center, Huntsville, AL, USA

SEP Mission to Titan NEXT Aerocapture In-Space Propulsion (Quicktime Movie)

Baggett, Randy; TECH ISP: Next Generation Ion; January 2004; In English; No Copyright; Avail: CASI; A01, Hardcopy

The ion thruster is one of the most promising solar electric propulsion (SEP) technologies to support future Outer Planet
missions (place provided link below here) for NASA's Office of Space Science. Typically, ion thrusters are used in high Isp-
low thrust applications that require long lifetimes, as well as, higher efficiency over state-of-the-art chemical propulsion
systems. Today, the standard for ion thrusters is the SEP Technology Application Readiness (NSTAR) thruster. Jet Propulsion
Laboratory's (JPL's) extended life test (ELT) of the DS 1 flight spare NSTAR thruster began in October 1998. This test
successfully demonstrated lifetime of the NSTAR flight spare thruster, which will provide a solid basis for selection of ion
thrusters for future Code S missions. The NSTAR ELT was concluded on June 30,2003 after 30,352 hours. The purpose of
the Next Generation Ion (NGI) activities is to advance Ion propulsion system technologies through the development of

10

NASA's Evolutionary Xenon Thruster (NEXT). The goal of NEXT is to more than double the power capability and lifetime
throughput (the total amount of propellant which can be processed) while increasing the Isp by 30% and the thrust by 120%.
Derived from text
Ion Propulsion; Solar Electric Propulsion

20030111896 Naval Postgraduate School, Monterey, CA
Optimization of Low Thrust Trajectories With Terminal Aerocapture

Josselyn, Scott B.; Jun. 2003; In English; Original contains color illustrations
Report No.(s): AD-A417512; No Copyright; Avail: CASI; A08, Hardcopy

This thesis explores using a direct pseudospectral method for the solution of optical control problems with mixed
dynamics. An easy to use MATLAB optimization package known as DIDO is used to obtain the solutions. The modeling of
both low thrust interplanetary trajectories as well as aerocapture trajectories is detailed and the solutions for low thrust
minimum time and minimum fuel trajectories are explored with particular emphasis on verification of the optimality of the
obtained solution. Optimal aerocapture trajectories are solved for rotating atmospheres over a range of arrival V- infinities.
Solutions are obtained using various performance indexes including minimum fuel, minimum heat load, and minimum total
aerocapture mass. Finally, the problem formulation and solutions for the mixed dynamic problem of low thrust trajectories
with a terminal aerocapture maneuver is addressed yielding new trajectories maximizing the total scientific mass at arrival.
This thesis explores using a direct pseudospectral method for the solution of optimal control problems with mixed dynamics.
An easy to use MATLAB optimization package known as DIDO is used to obtain the solutions. The modeling of both low
thrust interplanetary trajectories as well as aerocapture trajectories is detailed and the solutions for low thrust minimum time
and minimum fuel trajectories are explored with particular emphasis on verification of the optimality of the obtained solution.
Optimal aerocapture trajectories are solved for rotating atmospheres over a range of arrival V-infmities. Solutions are obtained
using various performance indexes including minimum fuel, minimum heat load, and minimum total aerocapture mass.
Finally, the problem formulation and solutions for the mixed dynamic problem of low thrust trajectories with a terminal
aerocapture maneuver is addressed yielding new trajectories maximizing the total scientific mass at arrival.
DTIC
Interplanetary Trajectories; Interorbital Trajectories; Trajectory Optimization; Optimal Control

20030107097 Air Force Inst, of Tech., Wright-Patterson AFB, OH
Aerocapture Guidance Methods for High Energy Trajectories

Dicarlo, Jennifer L.; May 23, 2003; In English

Report No.(s): AD-A416545; AFIT-CI02-1191; No Copyright; Avail: CASI; A07, Hardcopy

This thesis investigates enhancements of an existing numerical predictor-corrector aerocapture guidance algorithm
(PredGuid). The study includes implementation of an energy management phase prior to targeting with a generic method of
transition and replacement of heuristic features with more generic features. The vehicle response during energy management
was modeled as a second-order spring/mass/damper system. Phase change occurred when two conditions were met: First, the
vehicle could fly a constant bank angle of 1100 for the remainder of the trajectory and have the resulting apogee below or
within a given tolerance above the target apogee. Second, the predicted final energy indicated that the vehicle would be on
an elliptical, not hyperbolic, trajectory. So as to incorporate generic features, modeling of a separate lift down phase was
replaced by using a lift-down condition to determine phase change and biasing to the same lift- down condition during
targeting. Also, use of a heuristic sensitivity to calculate the first corrected bank angle was replaced by a simple smart
guessing' algorithm. Finally, heuristic lateral corridor boundaries were replaced by boundaries based on percentage of forward
velocity.
DTIC
Trajectories; Hyperbolic Trajectories; Predictor-Corrector Methods; Algorithms

20030106653 NASA Marshall Space Flight Center, Huntsville, AL, USA
Aerocapture Technology Project Overview

James, Bonnie; Munk, Michelle; Moon, Steve; July 20, 2003; In English

Report No.(s): AIAA Paper 2003-4654; No Copyright; Avail: CASI; A01, Hardcopy

Aerocapture technology development is one of the highest priority investments for the NASA In-Space Propulsion
Program (ISP). The ISP is managed by the NASA Headquarters Office of Space Science, and implemented by the Marshall
Space Flight Center in Huntsville, Alabama. The objective of the ISP Aerocapture Technology Project (ATP) is to develop

11

technologies that can enable and/or benefit NASA science missions by significantly reducing cost, mass, and trip times. To

accomplish this objective, the ATP identifies and prioritizes the most promising technologies using systems analysis,

technology advancement and peer review, coupled with NASA Headquarters Office of Space Science target requirements.

Efforts are focused on developing mid-Technology Readiness Level (TRL) technologies to systems-level spaceflight

validation.

Author

Aerocapture; Systems Analysis; Spacecraft Propulsion

20030106138 Ball Aerospace and Technologies Corp., Boulder, CO, USA
Trailing Ballute Aerocapture: Concept and Feasibility Assessment

Miller, Kevin L.; Gulick, Doug; Lewis, Jake; Trochman, Bill; Stein, Jim; Lyons, Daniel T; Wilmoth, Richard G.; July 21,
2003; In English; AIAA Joint Propulsion Conference and Exhibit 2003, 20-23 Jul. 2003, Huntsville, AL, USA
Contract(s)/Grant(s): NAS8-02130; JPL-1205966
Report No.(s): AIAA Paper 2003-4655; Copyright; Avail: CASI; A03, Hardcopy

Trailing Ballute Aerocapture offers the potential to obtain orbit insertion around a planetary body at a fraction of the mass
of traditional methods. This allows for lower costs for launch, faster flight times and additional mass available for science
payloads. The technique involves an inflated ballute (balloon-parachute) that provides aerodynamic drag area for use in the
atmosphere of a planetary body to provide for orbit insertion in a relatively benign heating environment. To account for
atmospheric, navigation and other uncertainties, the ballute is oversized and detached once the desired velocity change (Delta
V) has been achieved. Analysis and trades have been performed for the purpose of assessing the feasibility of the technique
including aerophysics, material assessments, inflation system and deployment sequence and dynamics, configuration trades,
ballute separation and trajectory analysis. Outlined is the technology development required for advancing the technique to a
level that would allow it to be viable for use in space exploration missions.
Author
Ballutes; Aerocapture; Aerodynamic Drag; Feasibility Analysis

20030105420

Mars reconnaissance orbiter design approach for high-resolution surface imaging

Lee, S. W.; Skulsky, E. D.; Chapel, J.; Cwynar, D.; Gehling, R.; Delamere, A.; Advances in the Astronautical Sciences; 2003;
ISSN 0065-3438; Volume 113, p. 509-528; In English; Guidance and Control 2003: Advances in the Astronautical Sciences,
Feb. 5-9, 2003, Breckenridge, CO, USA; Copyright; Avail: Other Sources

The Mars Reconnaissance Orbiter (MRO) will explore Mars equipped with a suite of six scientific instruments and two
engineering experiments, and supporting two additional facility investigations. One of the objectives of the MRO mission is
to use the High-Resolution Imaging Science Experiment (HiRISE) to provide 30 cm/pixel images of future Mars landing sites.
To achieve such detail, MRO must meet some very challenging target-relative pointing and pointing stability requirements.
A combination of analysis, operational constraints, and spacecraft design modifications were utilized to ensure that the
necessary pointing requirements will be met.
EI
High Resolution; Imaging Techniques; Orbits; Reconnaissance Aircraft

20030091868

Pitch control during autonomous aerobraking for near-term Mars exploration

Johnson, Wyatt R.; Longuski, James M.; Lyons, Daniel T; Journal of Spacecraft and Rockets; May/June 2003; ISSN
0022-4650; Volume 40, Issue no. 3, p. 371-379; In English; Copyright; Avail: Other Sources

Conventional aerobraking requires propellant to dump the spacecraft's angular momentum and to maintain attitude
control during the atmospheric fly through. We consider how reaction wheels can be used to control the spacecraft's pitch
during each atmospheric flythrough and to reduce angular momentum simultaneously. Control laws are developed for
minimum onboard instrumentation (where the only state information are the angular rates of the spacecraft and the reaction
wheels) to compensate for large variations in entry time and atmospheric density. Simulations indicate that pitch attitude and
angular momentum can be controlled with reaction wheels alone, thus saving precious propellant while significantly
increasing the timing margin for sequencing.
EI
Aerodynamics; Computerized Simulation; Drag; Spacecraft

12

20030080878

AIMS: Acousto-optic imaging spectrometer for spectral mapping of solid surfaces

Glenar, David A.; Blaney, Diana L.; Hillman, John J.; Acta Astronautica; January/March 2003; ISSN 0094-5765; Volume 52,
Issue no. 2-6, p. 389-396; In English; Copyright; Avail: Other Sources

A compact, two-channel acousto-optic tunable filter (AOTF) camera is being built at GSFC as a candidate payload
instrument for future Mars landers or small-body rendezvous missions. This effort is supported by the NASA Mars Instrument
Development Program (MIDP), Office of Space Science Advanced Technologies and Mission Studies. Acousto-optic Imaging
Spectrometer (AIMS) is electronically programmable and provides arbitrary spatial and spectral selection from 0.48 to 2.4 mu
m. The geometric throughput of AOTF' s are well matched to the requirements for lander mounted cameras since (I) they can
be made very compact, (II) 'slow' (f/14-f/18) optics required for large depth-of-field fall well within the angular aperture limit
of AOTF' s, and (III) they operate at low ambient temperatures. A breadboard of the AIMS short-wavelength channel is now
being used for spectral imaging of high-interest Mars analog materials (iron oxides, carbonates, sulfates and sedimentary
basalts) as part of the initial instrument validation exercises, (copyright) 2002 Published by Elsevier Science Ltd.
EI
Acousto-Optics; Aerospace Sciences; Cameras; Imaging Techniques; Planetary Landing

20030080863
Europa Lander

Gershman, Robert; Nilsen, Erik; Oberto, Robert; Acta Astronautica; January/March 2003; ISSN 0094-5765; Volume 52, Issue
no. 2-6, p. 253-258; In English; Copyright; Avail: Other Sources

A Europa Lander mission has been assigned high priority for the post-2005 time frame in NASA's Space Science
Enterprise Strategic Plan. Europa is one of the most scientifically interesting objects in the solar system because of the strong
possibility that a liquid water ocean exists underneath its ice-covered surface. The primary scientific goals of the proposed
Europa Lander mission are to characterize the surface material from a recent outflow and look for evidence of pre-biotic and
possibly biotic chemistry. The baseline mission concept involves landing a single spacecraft on the surface of Europa with the
capability to acquire samples of material, perform detailed chemical analysis of the samples, and transmit the results to Earth.
This paper provides a discussion of the benefits and status of the key spacecraft and instrument technologies needed to
accomplish the science objectives. Also described are variations on the baseline concept including the addition of small
auxiliary probes and an experimental ice penetration probe, (copyright) 2002 Elsevier Science Ltd. All rights reserved.
EI
Aerospace Sciences; Planetary Landing; Solar System; Spacecraft

20030066242 NASA Marshall Space Flight Center, Huntsville, AL, USA
Engineering-Level Model Atmospheres for Titan & Neptune

Justus, C. G.; Johnson, D. L.; July 20, 2003; In English, 20-23 Jul. 2003, Huntsville, AL, USA
Contract(s)/Grant(s): NAS8-60000; No Copyright; Avail: CASI; A01, Hardcopy

Engineering-level atmospheric models for Titan and Neptune have been developed for use in NASA s systems analysis
studies of aerocapture applications in missions to the outer planets. Analogous to highly successful Global Reference
Atmospheric Models for Earth (GRAM, Justus et al., 2000) and Mars (Mars-GRAM, Justus and Johnson, 2001, Justus et al.,
2002) the new models are called Titan-GRAM and Neptune-GRAM. Like GRAM and Mars-GRAM, an important feature of
Titan-GRAM and Neptune-GRAM is their ability to simulate quasi-random perturbations for Monte- Carlo analyses in
developing guidance, navigation and control algorithms, and for thermal systems design.
Author
Aerocapture; Titan; Monte Carlo Method; Neptune (Planet); Atmospheric Models

20030066102 Rhode Island Univ., Narragansett, RI, USA
Science and Engineering Potential of an Icy Moon Lander

DHondt, S. L.; Millerr, J. H.; Forum on Concepts and Approaches for Jupiter Icy Moons Orbiter; 2003, 17; In English;
Copyright; Abstract Only; Available from CASI only as part of the entire parent document

We urge consideration of an Icy Moon Lander as part of the Jupiter Icy Moon Orbiter mission. Inclusion of a lander would
have major advantages. It would allow scientific and engineering objectives to be met that cannot be addressed with an orbiter.
It would also allow independent tests of surface and subsurface properties inferred from remote observations. It would provide
invaluable engineering data for the design of a future ice or Ocean penetrator mission. We illustrate these advantages with

13

three examples. As the first example, an acoustic profiler imbedded in the surface of an icy moon could be used to identify
several subsurface properties as a function of depth. Some of these properties, such as the presence and depth (or absence)
of the water/ice interface and the structure and density of the ice as a function of depth, might be independently inferred by
instrumentation on an orbiter. Other properties that might be determinable with an acoustic profiler cannot be imaged from
orbit. These include the shear modulus of the ice (which might be used to distinguish between rigid ice and slushy convecting
ice), ocean density as a function of depth, the depth of an ocean/bedrock interface, and properties of the bedrock underlying
the ocean and ice. For the second example, instrumentation on a lander could undertake direct chemical analysis of organic
and inorganic compounds in the surface ice and atmosphere of an icy moon. Such analyses would directly test models of
surface compositions and atmospheric composition inferred from remote observations. These analyses would also greatly
advance human understanding of the chemical habitability of a Jovian icy moon by directly identifying and quantifying
concentrations of nutrients, energy yielding chemicals, and carbon sources in the surface ice. Thermal studies provide the third
example. Lander-based thermal measurements on the surface of an icy moon would provide an absolute calibration standard
for surface temperatures inferred from remote observations. Downhole temperature measurements taken at a single site with
a shallow penetrator would allow modeling of the subsurface temperature profile and independent estimation of the presence
and depth (or absence) of the ice/ocean interface. In closing, we wish to emphasize that inclusion of a lander with relatively
low-weight instrumentation in the JIMO mission would provide a high scientific pay-off. Because the lander instrumentation
would not penetrate the ice deeply, there would be no risk of directly contaminating any underlying ocean. Such a lander might
require only modest adaptation of existing technology and consequently might entail relatively low cost.
Derived from text
Galilean Satellites; Satellite Surfaces; Surface Properties; Planetary Landing; Measuring Instruments; Space Probes

20030065170 NASA Langley Research Center, Hampton, VA, USA

Mars Exploration Rover Six-Degree-Of-Freedom Entry Trajectory Analysis

Desai, Prasun N.; Schoenenberger, Mark; Cheatwood, F. M.; [2003]; In English, 3-7 Aug. 2003, Big Sky, MT, USA
Report No.(s): AAS Paper 03-642; Copyright; Avail: CASI; A03, Hardcopy

The Mars Exploration Rover mission will be the next opportunity for surface exploration of Mars in January 2004. Two
rovers will be delivered to the surface of Mars using the same entry, descent, and landing scenario that was developed and
successfully implemented by Mars Pathfinder. This investigation describes the trajectory analysis that was performed for the
hypersonic portion of the MER entry. In this analysis, a six-degree-of-freedom trajectory simulation of the entry is performed
to determine the entry characteristics of the capsules. In addition, a Monte Carlo analysis is also performed to statistically
assess the robustness of the entry design to off-nominal conditions to assure that all entry requirements are satisfied. The results
show that the attitude at peak heating and parachute deployment are well within entry limits. In addition, the parachute
deployment dynamics pressure and Mach number are also well within the design requirements.
Author

Degrees of Freedom; Trajectory Analysis; Atmospheric Entry; Mars Roving Vehicles; Mars Exploration; NASA Space
Programs

20030062242 NASA Marshall Space Flight Center, Huntsville, AL, USA
NASA Development of Aerocapture Technologies

James, Bonnie; Munk, Michelle; Moon, Steve; [2003]; In English, 20-22 May 2003, Monterey, CA, USA; Copyright; Avail:
CASI; A01, Hardcopy

Aeroassist technology development is a vital part of the NASA ln-Space Propulsion Program (ISP), which is managed
by the NASA Headquarters Office of Space Science, and implemented by the Marshall Space Flight Center in Huntsville,
Alabama. Aeroassist is the general term given to various techniques to maneuver a space vehicle within an atmosphere, using
aerodynamic forces in lieu of propulsive fuel. Within the ISP, the current aeroassist technology development focus is
aerocapture. The objective of the ISP Aerocapture Technology Project (ATP) is to develop technologies that can enable and/or
benefit NASA science missions by significantly reducing cost, mass, and/or travel times. To accomplish this objective, the ATP
identifies and prioritizes the most promising technologies using systems analysis, technology advancement and peer review,
coupled with NASA Headquarters Office of Space Science target requirements. Plans are focused on developing
mid-Technology Readiness Level (TRL) technologies to TRL 6 (ready for technology demonstration in space).
Author
NASA Space Programs; Aeroassist; Aerodynamic Forces; Systems Analysis; Propulsion; Aerocapture

14

20030055137

Development of a Monte Carlo Mars-gram model for 2001 Mars Odyssey aerobraking simulations

Dwyer, Alicia M.; Tolson, Robert H.; Munk, Michelle M.; Tartabini, Paul V.; Journal of the Astronautical Sciences; April- June
2002; ISSN 0021-9142; Volume 50, Issue no. 2, p. 191-211; In English; Copyright

Atmospheric density data taken during the Mars Global Surveyor aerobraking mission (1997-1999) showed significant
variability over the altitude range (100-140 km) of interest for aerobraking. This paper presents the method by which Mars
Global Surveyor data were used to determine the statistical distribution of mean density and the amplitude and phase of
stationary atmospheric waves as a function of latitude. The combination of mean density and waves produced a good fit to
the observed data. Using this information, a model was developed to implement the variations into Monte Carlo simulations
for future missions to Mars, specifically the Mars Odyssey aerobraking mission (October, 2001-January, 2002). An example
of Monte Carlo results for the Mars Odyssey aerobraking mission is shown.
EI
Atmospheric Density; Computerized Simulation; Monte Carlo Method; Planets; Space Flight

20030055136

Approaches to autonomous aerobraking at Mars

Hanna, J. L.; Tolson, R. H.; Journal of the Astronautical Sciences; April-June 2002; ISSN 0021-9142; Volume 50, Issue no.
2, p. 173-189; In English; Copyright

Planetary atmospheric aerobraking will most likely be incorporated in every future Mars orbiting mission. Aerobraking
requires an intensive workload during operations. To provide safe and efficient aerobraking, both navigation and spacecraft
system teams must be extremely diligent in updating spacecraft sequences and performing periapsis raise or lower maneuvers
to maintain the required orbital energy reduction without exceeding the design limits of the spacecraft. Automating the process
with onboard measurements could significantly reduce the operational burden and, in addition, could reduce the potential for
human error. Two levels of automation are presented and validated using part of the Mars Global Surveyor aerobraking
sequence and a simulated Mars Odyssey sequence. The simplest method only provides the capability to update the onboard
sequence. This method uses onboard accelerometer measurements to estimate the change in orbital period during an
aerobraking pass and thereby estimates the beginning of the next aerobraking sequence. Evaluation of the method utilizing
MGS accelerometer data showed that the time of the next periapsis can be estimated to within 25% 3 sigma of the change
in the orbital period due to drag. The second approach provides complete onboard orbit propagation. A low-order gravity
model is proposed that is sufficient to provide periapsis altitude predictions to within 100-200 meters over three orbits.
Accelerometer measurements are used as part of the trajectory force model while the spacecraft is in the atmosphere.
EI
Accelerometers; Navigation; Planets; Space Flight; Spacecraft

20030015758 NASA Langley Research Center, Hampton, VA USA

Wake Closure Characteristics and Afterbody Heating on a Mars Sample Return Orbiter

Horvath, Thomas J.; Cheatwood, McNeil F.; Wilmoth, Richard G.; Alter, Stephen J.; [2002]; In English, 3-7 Feb. 2002,
Albuquerque, NM, USA; No Copyright; Avail: CASI; A03, Hardcopy

Aeroheating wind-tunnel tests were conducted on a 0.028 scale model of an orbiter concept considered for a possible Mars
sample return mission. The primary experimental objectives were to characterize hypersonic near wake closure and determine
if shear layer impingement would occur on the proposed orbiter afterbody at incidence angles necessary for a Martian
aerocapture maneuver. Global heat transfer mappings, surface streamline patterns, and shock shapes were obtained in the
NASA Langley 20-Inch Mach 6 Air and CF4 Tunnels for post-normal shock Reynolds numbers (based on forebody diameter)
ranging from 1,400 to 415,000, angles of attack ranging from -5 to 10 degrees at 0, 3, and 6 degree sideslip, and normal-shock
density ratios of 5 and 12. Laminar, transitional, and turbulent shear layer impingement on the cylindrical afterbody was
inferred from the measurements and resulted in a localized heating maximum that ranged from 40 to 75 percent of the
reference forebody stagnation point heating. Comparison of laminar heating prediction to experimental measurement along the
orbiter afterbody highlight grid alignment challenges associated with numerical simulation of three- dimensional separated
wake flows. Predicted values of a continuum breakdown parameter revealed significant regions of non-continuum flow
downstream of the flow separation at the MSRO shoulder and in the region of the reattachment shock on the afterbody. The
presence of these regions suggest that the Navier-Stokes predictions at the laminar wind-tunnel condition may encounter errors
in the numerical calculation of the wake shear layer development and impingement due to non-continuum effects.
Author

Mars Sample Return Missions; Aerodynamic Heating; Wind Tunnel Tests; Hypersonic Wakes; Impingement; Aerocapture;
Interplanetary Spacecraft; Flow Characteristics

15

20030014800 NASA Langley Research Center, Hampton, VA USA
Autonomous Aerobraking at Mars

Hanna, Jill L.; Tolson, Robert; Cianciolo, Alicia Dwyer; Dec, John; [2002]; In English, 22-25 Oct. 2002, Frascati, Italy;
Original contains color illustrations; Copyright; Avail: CASI; A02, Hardcopy; Distribution as joint owner in the copyright

Aerobraking has become a proven approach for orbital missions at Mars. A launch of a 1000 kg class spacecraft on a Delta
class booster saves 90% of the post-MOI fuel otherwise required to circularize the orbit. In 1997, Mars Global Surveyor
demonstrated the feasibility and Mars 2001 Odyssey completed a nearly trouble free aerobraking phase in January 2002. In
2006, Mars Reconnaissance Orbiter will also utilize aerobraking. From the flight operations standpoint, however, aerobraking
is labor intensive and high risk due to the large density variability in the Mars thermosphere. The maximum rate of aerobraking
is typically limited by the maximum allowable temperature of the solar array which is the primary drag surface. Prior missions
have used a surrogate variable, usually maximum free stream heat flux, as a basis for performing periapsis altitude corridor
control maneuvers. This paper provides an adaptive sequential method for operationally relating measured temperatures to
heat flux profile characteristics and performing maneuvers based directly on measured temperatures and atmospheric
properties derived from the heat flux profiles. Simulations of autonomous aerobraking are performed using Odyssey mission
data.
Author

Aerobraking; Mars Missions; Spacecraft Maneuvers; Aeromaneuvering; Flight Operations; Computerized Simulation;
Temperature Profiles; Solar Arrays; Heat Flux

20030014794 NASA Langley Research Center, Hampton, VA USA

Multibody Parachute Flight Simulations for Planetary Entry Trajectories Using 'Equilibrium Points'

Raiszadeh, Ben; [2003]; In English, 9-13 Feb. 2003, Ponce, Puerto Rico; Original contains color illustrations
Report No.(s): AAS-03-163; No Copyright; Avail: CASI; A03, Hardcopy

A method has been developed to reduce numerical stiffness and computer CPU requirements of high fidelity multibody
flight simulations involving parachutes for planetary entry trajectories. Typical parachute entry configurations consist of entry
bodies suspended from a parachute, connected by flexible lines. To accurately calculate line forces and moments, the
simulations need to keep track of the point where the flexible lines meet (confluence point). In previous multibody parachute
flight simulations, the confluence point has been modeled as a point mass. Using a point mass for the confluence point tends
to make the simulation numerically stiff, because its mass is typically much less that than the main rigid body masses. One
solution for stiff differential equations is to use a very small integration time step. However, this results in large computer CPU
requirements. In the method described in the paper, the need for using a mass as the confluence point has been eliminated.
Instead, the confluence point is modeled using an 'equilibrium point' . This point is calculated at every integration step as the
point at which sum of all line forces is zero (static equilibrium). The use of this 'equilibrium point' has the advantage of both
reducing the numerical stiffness of the simulations, and eliminating the dynamical equations associated with vibration of a
lumped mass on a high-tension string.
Author
Atmospheric Entry; Flight Simulation; Parachutes; Trajectories; Differential Equations; Computerized Simulation

20030014283 NASA Langley Research Center, Hampton, VA USA
Plume Modeling and Application to Mars 2001 Odyssey Aerobraking

Chavis, Zachary Q.; Wilmoth, Richard G.; [2002]; In English, 24-26 Jun. 2002, Saint Louis, MO, USA; Original contains
color illustrations; No Copyright; Avail: CASI; A03, Hardcopy

A modified source flow model was used to calculate the plume flowfield from a Mars Odyssey thruster during
aerobraking. The source flow model results compared well with previous detailed CFD results for a Mars Global Surveyor
thruster. Using an iso-density surface for the Odyssey plume, DSMC simulations were performed to determine the effect the
plumes have on the Odyssey aerodynamics. A database was then built to incorporate the plume effects into 6-DOF simulations
over a range of attitudes and densities expected during aerobraking. 6-DOF simulations that included the plume effects showed
better correlation with flight data than simulations without the plume effects.
Author

Computational Fluid Dynamics; Aerobraking; 2001 Mars Odyssey; Computerized Simulation; Rocket Exhaust; Flow
Distribution; Mathematical Models

16

20030006687 NASA Langley Research Center, Hampton, VA USA

Thermal Analysis and Correlation of the Mars Odyssey Spacecraft's Solar Array During Aerobraking Operations

Dec, John A.; Gasbarre, Joseph R; George, Benjamin E.; [2002]; In English, 5-8 Aug. 2002, Monterey, CA, USA; Original

contains color illustrations

Report No.(s): AIAA Paper 2002-4536; Copyright; Avail: CASI; A03, Hardcopy; Distribution under U.S. Government purpose

rights

The Mars Odyssey spacecraft made use of multipass aerobraking to gradually reduce its orbit period from a highly
elliptical insertion orbit to its final science orbit. Aerobraking operations provided an opportunity to apply advanced thermal
analysis techniques to predict the temperature of the spacecraft's solar array for each drag pass. Odyssey telemetry data was
used to correlate the thermal model. The thermal analysis was tightly coupled to the flight mechanics, aerodynamics, and
atmospheric modeling efforts being performed during operations. Specifically, the thermal analysis predictions required a
calculation of the spacecraft's velocity relative to the atmosphere, a prediction of the atmospheric density, and a prediction
of the heat transfer coefficients due to aerodynamic heating. Temperature correlations were performed by comparing predicted
temperatures of the thermocouples to the actual thermocouple readings from the spacecraft. Time histories of the spacecraft
relative velocity, atmospheric density, and heat transfer coefficients, calculated using flight accelerometer and quaternion data,
were used to calculate the aerodynamic heating. During aerobraking operations, the correlations were used to continually
update the thermal model, thus increasing confidence in the predictions. This paper describes the thermal analysis that was
performed and presents the correlations to the flight data.
Author
Thermal Analysis; Correlation; 2001 Mars Odyssey; Solar Arrays; Heat Transfer Coefficients; Aerodynamic Heating

20030006120 NASA Langley Research Center, Hampton, VA USA

Control Surface and Afterbody Experimental Aeroheating for a Proposed Mars Smart Lander Aeroshell

Liechty, Derek S.; Hollis, Brian R.; Edquist, Karl T; [2002]; In English, 5-8 Aug. 2002, Monterey, CA, USA; Original

contains color illustrations

Report No.(s): AIAA Paper 2002-4506; Copyright; Avail: CASI; A03, Hardcopy; Distribution under U.S. Government purpose

rights

Several configurations, having a Viking aeroshell heritage and providing lift-to-drag required for precision landing, have
been considered for a proposed Mars Smart Lander. An experimental aeroheating investigation of two configurations, one
having a blended tab and the other a blended shelf control surface, has been conducted at the NASA Langley Research Center
in the 20-Inch Mach 6 Air Tunnel to assess heating levels on these control surfaces and their effects on afterbody heating. The
proposed Mars Smart Lander concept is to be attached through its aeroshell to the main spacecraft bus, thereby producing
cavities in the forebody heat shield upon separation prior to entry into the Martian atmosphere. The effects these cavities will
have on the heating levels experienced by the control surface and the afterbody were also examined. The effects of Reynolds
number, angle-of-attack, and cavity location on aeroheating levels and distributions were determined and are presented. At the
highest angle-of-attack, blended tab heating was increased due to transitional reattachment of the separated shear layer. The
placement of cavities downstream of the control surface greatly influenced aeroheating levels and distributions. Forebody heat
shield cavities had no effect on afterbody heating and the presence of control surfaces decreased leeward afterbody heating
slightly.
Author
Control Surfaces; Wind Tunnel Tests; Aeroshells; Aerodynamic Heating; Mars Landing

20030005808 NASA Langley Research Center, Hampton, VA USA

The Development and Evaluation of an Operational Aerobraking Strategy for the Mars 2001 Odyssey Orbiter

Tartabini, Paul V; Munk, Michelle M.; Powell, Richard W.; [2002]; In English, 5-8 Aug. 2002, Monterey, CA, USA; Original
contains color illustrations

Report No.(s): AIAA Paper 2002-4537; No Copyright; Avail: CASI; A03, Hardcopy; Distribution under U.S. Government
purpose rights

The Mars 2001 Odyssey Orbiter successfully completed the aerobraking phase of its mission on January 11, 2002. This
paper discusses the support provided by NASA's Langley Research Center to the navigation team at the Jet Propulsion
Laboratory in the planning and operational support of Mars Odyssey Aerobraking. Specifically, the development of a
three-degree-of-freedom aerobraking trajectory simulation and its application to pre-flight planning activities as well as
operations is described. The importance of running the simulation in a Monte Carlo fashion to capture the effects of mission
and atmospheric uncertainties is demonstrated, and the utility of including predictive logic within the simulation that could

17

mimic operational maneuver decision-making is shown. A description is also provided of how the simulation was adapted to

support flight operations as both a validation and risk reduction tool and as a means of obtaining a statistical basis for

maneuver strategy decisions. This latter application was the first use of Monte Carlo trajectory analysis in an aerobraking

mission.

Author

Aerobraking; Capture Effect; Flight Operations; Navigation; Planning; Simulation; Trajectories

20030005452 NASA Johnson Space Center, Houston, TX USA
Aerocapture Guidance Algorithm Comparison Campaign

Rousseau, Stephane; Perot, Etienne; Graves, Claude; Masciarelli, James P.; Queen, Eric; [2002]; In English, 5-8 Aug. 2002,
Monterey, CA, USA; Original contains color illustrations

Report No.(s): AIAA Paper 2002-4822; Copyright; Avail: CASI; A02, Hardcopy; Distribution as joint owner in the copyright
The aerocapture is a promising technique for the future human interplanetary missions. The Mars Sample Return was
initially based on an insertion by aerocapture. A CNES orbiter Mars Premier was developed to demonstrate this concept.
Mainly due to budget constraints, the aerocapture was cancelled for the French orbiter. A lot of studies were achieved during
the three last years to develop and test different guidance algorithms (APC, EC, TPC, NPC). This work was shared between
CNES and NASA, with a fruitful joint working group. To finish this study an evaluation campaign has been performed to test
the different algorithms. The objective was to assess the robustness, accuracy, capability to limit the load, and the complexity
of each algorithm. A simulation campaign has been specified and performed by CNES, with a similar activity on the NASA
side to confirm the CNES results. This evaluation has demonstrated that the numerical guidance principal is not competitive
compared to the analytical concepts. All the other algorithms are well adapted to guaranty the success of the aerocapture. The
TPC appears to be the more robust, the APC the more accurate, and the EC appears to be a good compromise.
Author
Aerocapture; Spacecraft Guidance; Algorithms

20030005447 NASA Langley Research Center, Hampton, VA USA

Experimental Hypersonic Aerodynamic Characteristics of the 2001 Mars Surveyor Precision Lander with Flap

Horvath, Thomas J.; OConnell, Tod F; Cheatwood, F. McNeil; Prabhu, Ramadas K.; Alter, Stephen J.; [2002]; In English, 5-8
Aug. 2002, Monterey, CA, USA

Report No.(s): AIAA Paper 2002-4408; Copyright; Avail: CASI; A03, Hardcopy; Distribution as joint owner in the copyright
Aerodynamic wind-tunnel screening tests were conducted on a 0.029 scale model of a proposed Mars Surveyor 2001
Precision Lander (70 deg half angle spherically blunted cone with a conical afterbody). The primary experimental objective
was to determine the effectiveness of a single flap to trim the vehicle at incidence during a lifting hypersonic planetary entry.
The laminar force and moment data, presented in the form of coefficients, and shock patterns from schlieren photography were
obtained in the NASA Langley Aerothermodynamic Laboratory for post-normal shock Reynolds numbers (based on forebody
diameter) ranging from 2,637 to 92,350, angles of attack ranging from tip to 23 degrees at and 2 degree sideslip, and
normal-shock density ratios of 5 and 12. Based upon the proposed entry trajectory of the 2001 Lander, the blunt body heavy
gas tests in CF, simulate a Mach number of approximately 12 based upon a normal shock density ratio of 12 in flight at Mars.
The results from this experimental study suggest that when traditional means of providing aerodynamic trim for this class of
planetary entry vehicle are not possible (e.g. offset e.g.), a single flap can provide similar aerodynamic performance. An
assessment of blunt body aerodynamic effects attributed to a real gas were obtained by synergistic testing in Mach 6 ideal-air
at a comparable Reynolds number. From an aerodynamic perspective, an appropriately sized flap was found to provide
sufficient trim capability at the desired L/D for precision landing. Inviscid hypersonic flow computations using an unstructured
grid were made to provide a quick assessment of the Lander aerodynamics. Navier-Stokes computational predictions were
found to be in very good agreement with experimental measurement.
Author
Aerodynamic Characteristics; Aerothermodynamics; Hypersonic Flow; Inviscid Flow

20030002240 NASA Langley Research Center, Hampton, VA USA
Computational Analysis of Towed Ballute Interactions

Gnoffo, Peter A.; Anderson, Brian P.; [2002]; In English, 24-26 Jun. 2002, Saint Louis, MO, USA; Original contains color

illustrations

Report No.(s): AIAA Paper 2002-2997; Copyright; Avail: CASI; A03, Hardcopy; Distribution as joint owner in the copyright

18

A ballute (balloon-parachute) is an inflatable, aerodynamic drag device for application to planetary entry vehicles.
Ballutes may be directly attached to a vehicle, increasing its cross-sectional area upon inflation, or towed behind the vehicle
as a semi-independent device that can be quickly cut free when the requisite change in velocity is achieved. The
aero thermodynamics of spherical and toroidal towed ballutes are considered in the present study. A limiting case of zero
towline length (clamped system) is also considered. A toroidal system can be designed (ignoring influence of the tethers) such
that all flow processed by the bow shock of the towing spacecraft passes through the hole in the toroid. For a spherical ballute,
towline length is a critical parameter that affects aeroheating on the ballute being towed through the spacecraft wake. In both
cases, complex and often unsteady interactions ensue in which the spacecraft and its wake resemble an aero spike situated in
front of the ballute. The strength of the interactions depends upon system geometry and Reynolds number. We show how
interactions may envelope the base of the towing spacecraft or impinge on the ballute surface with adverse consequences to
its thermal protection system. Geometric constraints to minimize or eliminate such adverse interactions are discussed. The
towed, toroidal system and the clamped, spherical system show greatest potential for a baseline design approach.
Author
Atmospheric Entry; Ballutes; Spacecraft Control; Flight Control; Towed Bodies; Computerized Simulation

20030002226 NASA Langley Research Center, Hampton, VA USA

Application of Accelerometer Data to Mars Odyssey Aerobraking and Atmospheric Modeling

Tolson, R. H.; Keating, G. M.; George, B. E.; Escalera, P. E.; Werner, M. R.; Dwyer, A. M.; Hanna, J. L.; [2002]; In English,
5-8 Aug. 2002, Monterey, CA, USA; Original contains color illustrations

Report No.(s): AIAA Paper 2002-4533; Copyright; Avail: CASI; A03, Hardcopy; Distribution as joint owner in the copyright
Aerobraking was an enabling technology for the Mars Odyssey mission even though it involved risk due primarily to the
variability of the Mars upper atmosphere. Consequently, numerous analyses based on various data types were performed
during operations to reduce these risk and among these data were measurements from spacecraft accelerometers. This paper
reports on the use of accelerometer data for determining atmospheric density during Odyssey aerobraking operations.
Acceleration was measured along three orthogonal axes, although only data from the component along the axis nominally into
the flow was used during operations. For a one second count time, the RMS noise level varied from 0.07 to 0.5 mm/s2
permitting density recovery to between 0.15 and 1.1 kg per cu km or about 2% of the mean density at periapsis during
aerobraking. Accelerometer data were analyzed in near real time to provide estimates of density at periapsis, maximum
density, density scale height, latitudinal gradient, longitudinal wave variations and location of the polar vortex. Summaries are
given of the aerobraking phase of the mission, the accelerometer data analysis methods and operational procedures, some
applications to determining thermospheric properties, and some remaining issues on interpretation of the data. Pre-flight
estimates of natural variability based on Mars Global Surveyor accelerometer measurements proved reliable in the
mid-latitudes, but overestimated the variability inside the polar vortex.
Author

Accelerometers; 2001 Mars Odyssey; Aerobraking; Atmospheric Models; Mars Atmosphere; Atmospheric Density; Radio
Tracking; Numerical Analysis; Trajectory Analysis

20030000829 NASA Langley Research Center, Hampton, VA USA
Mars Smart Lander Parachute Simulation Model

Queen, Eric M.; Raiszadeh, Ben; [2002]; In English, 5-8 Aug. 2002, Monterey, CA, USA

Report No.(s): AIAA Paper 2002-4617; Copyright; Avail: CASI; A02, Hardcopy; Distribution under U.S. Government purpose

rights

A multi-body flight simulation for the Mars Smart Lander has been developed that includes six degree-of-freedom
rigid-body models for both the supersonically-deployed and subsonically-deployed parachutes. This simulation is designed to
be incorporated into a larger simulation of the entire entry, descent and landing (EDL) sequence. The complete end-to-end
simulation will provide attitude history predictions of all bodies throughout the flight as well as loads on each of the connecting
lines. Other issues such as recontact with jettisoned elements (heat shield, back shield, parachute mortar covers, etc.), design
of parachute and attachment points, and desirable line properties can also be addressed readily using this simulation.
Author
Flight Simulation; Parachutes; Mars Landing; Trajectory Analysis; Evaluation; Loads (Forces)

20030000735 CFD Research Corp., Huntsville, AL USA

CFD Prediction of the BEAGLE 2 Mars Probe Aerodynamic Database

Liever, Peter A.; Habchi, Sami D.; Burnell, Simon I.; Lingard, Steve J.; Twelfth Thermal and Fluids Analysis Workshop; July
2002; In English; Original contains color illustrations; No Copyright; Avail: CASI; A03, Hardcopy

19

CFD (Computational Fluid Dynamics) has matured to provide reliable planetary entry vehicle aerodynamic predictions.
CFD provides substantial time and cost savings. CFD-FASTRAN was applied over the entire trajectory (Entry to Chute
Deployment). It provided valuable insight into vehicle flow characteristics (Examples: Wake and Base Flow Structure,
Transonic Wake Unsteadiness). A blended aerodynamic database was generated by combining CFD data, scaled existing data,
and wind tunnel test data. CFD based pitch damping analysis provides insight into dynamic stability characteristics not easily
obtained from wind tunnel tests.
Derived from text
Computational Fluid Dynamics; Wind Tunnel Tests; Atmospheric Entry; Predictions; Data Bases

20020081342 NASA Glenn Research Center, Cleveland, OH USA
Radioisotope Electric Propulsion for Fast Outer Planetary Orbiters

Oleson, Steven; Benson, Scott; Gefert, Leon; Patterson, Michael; Schreiber, Jeffrey; September 2002; In English, 7-10 Jul.

2002, Indianapolis, IN, USA

Contract(s)/Grant(s): RTOP 344-96-8D

Report No.(s): NASA/TM-2002-211893; NAS 1.15:211893; E-13575; AIAA Paper 2002-3967; No Copyright; Avail: CASI;

A03, Hardcopy

Recent interest in outer planetary targets by the Office of Space Science has spurred the search for technology options to
enable relatively quick missions to outer planetary targets. Several options are being explored including solar electric propelled
stages combined with aerocapture at the target and nuclear electric propulsion. Another option uses radioisotope powered
electric thrusters to reach the outer planets. Past work looked at using this technology to provide faster fiybys. A better use
for this technology is for outer planet orbiters. Combined with medium class launch vehicles and a new direct trajectory these
small, sub-kilowatt ion thrusters and Stirling radioisotope generators were found to allow missions as fast as 5 to 12 years for
objects from Saturn to Pluto, respectively. Key to the development is light spacecraft and science payload technologies.
Author
Nuclear Electric Propulsion; Radioactive Isotopes; Gas Giant Planets; Grand Tours; Aerocapture

20020077966 NASA Ames Research Center, Moffett Field, CA USA
Aerocapture Technology Development Needs for Outer Planet Exploration

Wercinski, Paul; Munk, Michelle; Powell, Richard; Hall, Jeff; Graves, Claude; Partridge, Harry, Technical Monitor; January

2002; In English

Contract(s)/Grant(s): RTOP 713-81-80

Report No.(s): NASA/TM-2002-211386; NAS 1.15:211386; A-0107378; No Copyright; Avail: CASI; A03, Hardcopy

The purpose of this white paper is to identify aerocapture technology and system level development needs to enable
NASA future mission planning to support Outer Planet Exploration. Aerocapture is a flight maneuver that takes place at very
high speeds within a planet's atmosphere that provides a change in velocity using aerodynamic forces (in contrast to
propulsive thrust) for orbit insertion. Aerocapture is very much a system level technology where individual disciplines such
as system analysis and integrated vehicle design, aerodynamics, aerothermal environments, thermal protection systems (TPS),
guidance, navigation and control (GN&C) instrumentation need to be integrated and optimized to meet mission specific
requirements. This paper identifies on-going activities, their relevance and potential benefit to outer planet aerocapture that
include New Millennium ST7 Aerocapture concept definition study, Mars Exploration Program aeroassist project level
support, and FY01 Aeroassist In-Space Guideline tasks. The challenges of performing aerocapture for outer planet missions
such as Titan Explorer or Neptune Orbiter require investments to advance the technology readiness of the aerocapture
technology disciplines for the unique application of outer planet aerocapture. This white paper will identify critical technology
gaps (with emphasis on aeroshell concepts) and strategies for advancement.
Author
Aerocapture; Aerothermodynamics; Spacecraft Design; Orbit Insertion; Spacecraft Maneuvers; Outer Planets Explorers

20020039836 NASA Langley Research Center, Hampton, VA USA

Aerothermal Instrumentation Loads To Implement Aeroassist Technology in Future Robotic and Human Missions to

MARS and Other Locations Within the Solar System

Parmar, Devendra S.; Shams, Qamar A.; April 2002; In English

Contract(s)/Grant(s): RTOP 713-81-70

Report No.(s): NASA/TM-2002-211459; NAS 1.15:211459; L-18123; No Copyright; Avail: CASI; A03, Hardcopy

20

The strategy of NASA to explore space objects in the vicinity of Earth and other planets of the solar system includes
robotic and human missions. This strategy requires a road map for technology development that will support the robotic
exploration and provide safety for the humans traveling to other celestial bodies. Aeroassist is one of the key elements of
technology planning for the success of future robot and human exploration missions to other celestial bodies. Measurement
of aerothermodynamic parameters such as temperature, pressure, and acceleration is of prime importance for aeroassist
technology implementation and for the safety and affordability of the mission. Instrumentation and methods to measure such
parameters have been reviewed in this report in view of past practices, current commercial availability of instrumentation
technology, and the prospects of improvement and upgrade according to the requirements. Analysis of the usability of each
identified instruments in terms of cost for efficient weight-volume ratio, power requirement, accuracy, sample rates, and other
appropriate metrics such as harsh environment survivability has been reported.
Author

Aeroassist; Aerothermodynamics; Robotics; Technology Utilization; Aerodynamic Loads; Solar System; Manned Mars
Missions; Temperature Measuring Instruments

20020023456 Instituto Nacional de Pesquisas Espacias, Sao Jose dos Campos, Brazil
Study of Orbital Transfers with Aeroassisted Maneuvers

Schulz, Walkiria; 2001; In Portuguese

Report No.(s): INPE-8476-TDI/776; No Copyright; Avail: CASI; A09, Hardcopy

This work presents an analysis of space missions through the development of a software package for the calculation of
aerodynamic maneuvers and of the required thrust maneuvers for their implementation. Besides the numerical development,
an analytical study contemplates the accomplishment of the aeroassisted phase of this maneuver type. This analysis includes
a study of the thermal limits associated with a vehicle passage through the atmosphere as well as an analysis of the associated
errors. Several simulations of aerodynamic maneuvers are carried out and compared with orbital changes accomplished
outside of the atmosphere. Among the conclusions, it is shown that the problem is extremely sensitive to the initial conditions
and each mission deserves a careful individual analysis. Finally, the results obtained from the analytical formulation are shown
to be in agreement with the numerical results for the upper layers of the terrestrial atmosphere.
Author
Transfer Orbits; Space Missions; Applications Programs (Computers); Aerodynamic Characteristics; Aeroassist

20020002105 NASA Johnson Space Center, Houston, TX USA
Beagle 2: The Next Exobiology Mission to Mars

Gibson, Everett K., Jr.; Pillinger, Colin T; Wright, Ian P.; Morse, Andy; Stewart, Jenny; Morgan, G.; Praine, Ian; Leigh,
Dennis; Sims, Mark R.; General Meeting of the NASA Astrobiology Insititute; April 2001, 160-162; In English; No
Copyright; Avail: Other Sources; Abstract Only

Beagle 2 is a 60 kg probe (with a 30 kg lander) developed in the UK for inclusion on the European Space Agency's 2003
Mars Express. Beagle 2 will deliver to the Martian surface a payload which consists of a high percentage of science
instruments to landed spacecraft mass. Beagle 2 will be launched in June, 2003 with Mars Express on a Soyuz-Fregat rocket
from the Baikonur Cosmodrome in Kazakhstan. Beagle 2 will land on Mars on December 26, 2003 in the Isidis Planitia basin
(approximately 10 degrees N and 275 degrees W), a large sedimentary basin that overlies the boundary between ancient
highlands and northern plains. Isidis Planitia, the third largest basin on Mars, which is possibly filled with sediment deposited
at the bottom of long-standing lakes or seas, offers an ideal environment for preserving traces of life. Beagle 2 was developed
to search for organic material and other volatiles on and below the surface of Mars in addition to the study of the inorganic
chemistry and mineralogy. Beagle 2 will utilize a mechanical mole and grinder to obtain samples from below the surface,
under rocks and inside rocks. A pair of stereo cameras will image the landing site along with a microscope for examination
of surface and rock samples. Analyses will include both rock and soil samples at various wavelengths, X-ray spectrometer and
Mossbauer spectrometer as well as a search for organics and other light element species (e.g. carbonates and water) and
measurement of their isotopic compositions. Beagle 2 has as its focus the goal of establishing whether evidence for life existed
in the past on Mars at the Isidis Planitia site or at least establishing if the conditions were ever suitable. Carbonates and organic
components were first recognized as existing on Mars when they were found in the Martian meteorite Nakhla. Romanek et
al showed the carbonates in ALH84001 were formed at low temperatures. McKay et al noted possible evidence of early life
on Mars within the ALH84001 meteorite. Thomas-Keprta et al showed the magnetite biomarkers in ALH84001's carbonates
are indistinguishable from those formed by magnetotactic bacteria found on Earth. Gibson et al showed there was significant
evidence for liquid water and biogenic products present on Mars across a 3.9 billion year period. A mechanical arm (PAW)
operates from the lander and is used for science operations along with sample acquisition. Instruments attached to the PAW

21

include: stereo cameras, Mossbauer instrument, X-ray fluorescence instrument, microscope, environmental sensors, rock
corer/grinder, a spoon, mirror, brushes, a mole attachment for acquisition of subsurface to depths of 1 to 2 meters and an
illumination device. Each camera has 14 filters which have been optimized for mineralogy composition, dust and water vapor
detection. The microscope's camera is designed for viewing the size and shape of dust particles, rock surfaces, microfossils,
and characteristics of the samples prior to introduction into the gas analysis package (GAP). The camera has a resolution of
4 microns/pixel, features four color capability (red, green, blue and UV (ultraviolet) fluoresence), a depth of focus of 40
micrometers and translation stage of +3 millimeters. The heart of the Beagle 2's life detection package is the gas analysis
package (GAP), which consists of a mass spectrometer with collectors at fixed masses for precise isotopic ratio measurements
and voltage scanning for spectral analysis. Primary aim of the GAP is to search for the presence of bulk constituents, individual
species, and isotopic fractionations for both extinct and extent life along with studying the low-temperature geochemistry of
the hydrogen, carbon, nitrogen and oxygen components on Mars from both the surface and atmosphere. GAP is a magnetic
sector mass spectrometer with the range of 1 to 140 amu which can be operated in both the static and dynamic modes. A triple
Faraday collector array will be used for C, N and O ratios along with a double Faraday array for H/D. Pulse counting electron
multiplier will be utilized for noble gases and selected organics. Anticipated detection limits are at the picomole level for
operation in the static mode of operation and high precision isotopic measurements will be made in the dynamic mode. Sample
processing and preparation system consists of reaction vessels along with references. Sample ovens capable of being heated
are attached to the manifold for sample combustion. Surface, subsurface materials and interior rock specimens will be
combusted in pure oxygen gas at various temperature intervals to release organic matter and volatiles. Combustion process
will permit detection of all forms and all atoms of carbon present in the samples. A chemical processing system is capable of
a variety of conversion reactions. Gases are manipulated either by cryogenic or chemical reactions and passed through the gas
handling portion of the vacuum system. There are two modes of operation: quantitative analysis and precise isotopic
measurements. Additional information is contained in the original extended abstract.
Author
Exobiology; Mars Missions; Spacecraft Instruments; European Space Programs; Space Probes

20010122748 Tennessee Univ., Knoxville, TN USA

Remote Sensing of Evaporite Minerals in Badwater Basin, Death Valley, at Varying Spatial Scales and in Different

Spectral Regions

Moersch, J. E.; Farmer, J.; Baldridge, A.; Field Trip and Workshop on the Martian Highlands and Mojave Desert Analogs;
2001, 45-46; In English; No Copyright; Avail: CASI; A01, Hardcopy

A key concept behind the overall architecture of NASA's Mars Surveyor Program is that remote sensing observations
made from orbit will be used to guide the selection of landing sites for subsequent missions to the surface. An important
component of the orbital phase of this strategy is mineralogical mapping of the surface with infrared spectrometers and
imaging systems. Currently, the Mars Global Surveyor Thermal Emission Spectrometer (TES) is spectrally mapping Mars in
the 6-50 micrometer region at a spatial resolution of 3 km. Starting later this year, the Thermal Emission Imaging System
(THEMIS) aboard the Mars 2001 Odyssey orbiter will image the entire surface of the planet in eight broad bands in the
6.5-14.5 micrometer region at a spatial resolution of 100 m. In 2003, ESA plans to launch the OMEGA instrument on Mars
Express, which will map the planet in the visible and near infrared regions from an elliptical orbit at spatial resolutions of up
to 100 m. Currently, NASA is selecting a visible and near-infrared mapping spectrometer for an orbiter that will launch in
2005. This instrument will likely map at a constant spatial resolution of at least 50 m. From an astrobiological perspective,
the utility of these spectral datasets will be in locating potential paleohabitats for martian life, via the detection of minerals
that form in the presence of liquid water. Deposits of evaporite minerals in putative martian paleolake basins are a particularly
attractive target to look for because of the areal extent of these features, the strong spectral features of these minerals, and the
characteristic sequences in which they appear along the margin of a basin. Despite considerable geomorphic evidence
indicating the presence of ancient lake basins on Mars, to date no evaporite deposits have been reported from the TES
experiment. But is this to be expected, given the limited spatial resolution of TES data? Might we still hope to find such
deposits in upcoming experiments? One way to address this question is to use existing datasets from terrestrial analog sites
to attempt to determine spatial and spectral thresholds of detectability for these minerals in a natural setting
Author

Remote Sensing; Death Valley (CA); Spatial Resolution; Spectral Bands; Mineral Deposits; Structural Basins; Sedimentary
Rocks

22

20010099686 Georgia Inst, of Tech., Atlanta, GA USA

Uncertainty Optimization Applied to the Monte Carlo Analysis of Planetary Entry Trajectories

Olds, John; Way, David; Jul. 31, 2001; In English

Contract(s)/Grant(s): NGT1-52163; No Copyright; Avail: CASI; All, Hardcopy

Recently, strong evidence of liquid water under the surface of Mars and a meteorite that might contain ancient microbes
have renewed interest in Mars exploration. With this renewed interest, NASA plans to send spacecraft to Mars approx. every
26 months. These future spacecraft will return higher-resolution images, make precision landings, engage in longer-ranging
surface maneuvers, and even return Martian soil and rock samples to Earth. Future robotic missions and any human missions
to Mars will require precise entries to ensure safe landings near science objective and pre-employed assets. Potential sources
of water and other interesting geographic features are often located near hazards, such as within craters or along canyon walls.
In order for more accurate landings to be made, spacecraft entering the Martian atmosphere need to use lift to actively control
the entry. This active guidance results in much smaller landing footprints. Planning for these missions will depend heavily on
Monte Carlo analysis. Monte Carlo trajectory simulations have been used with a high degree of success in recent planetary
exploration missions. These analyses ascertain the impact of off-nominal conditions during a flight and account for uncertainty.
Uncertainties generally stem from limitations in manufacturing tolerances, measurement capabilities, analysis accuracies, and
environmental unknowns. Thousands of off-nominal trajectories are simulated by randomly dispersing uncertainty variables
and collecting statistics on forecast variables. The dependability of Monte Carlo forecasts, however, is limited by the accuracy
and completeness of the assumed uncertainties. This is because Monte Carlo analysis is a forward driven problem; beginning
with the input uncertainties and proceeding to the forecasts outputs. It lacks a mechanism to affect or alter the uncertainties
based on the forecast results. If the results are unacceptable, the current practice is to use an iterative, trial-and-error approach
to reconcile discrepancies. Therefore, an improvement to the Monte Carlo analysis is needed that will allow the problem to
be worked in reverse. In this way, the largest allowable dispersions that achieve the required mission objectives can be
determined quantitatively.
Derived from text

Atmospheric Entry; Trajectory Optimization; Monte Carlo Method; Trajectory Planning; Trajectory Analysis; Trajectory
Control; Mathematical Models

20010041296 NASA Ames Research Center, Moffett Field, CA USA
Exploration of Titan Using Vertical Lift Aerial Vehicles

Young, L. A.; Forum on Innovative Approaches to Outer Planetary Exploration 2001-2020; 2001, 94; In English; No
Copyright; Abstract Only; Available from CASI only as part of the entire parent document

Autonomous vertical lift aerial vehicles (such as rotorcraft or powered-lift vehicles) hold considerable potential for
supporting planetary science and exploration missions. Vertical lift aerial vehicles would have the following advantages/
attributes for planetary exploration: (1) low-speed and low-altitude detailed aerial surveys; (2) remote-site sample return to
lander platforms; (3) precision placement of scientific probes; (4) soft landing capability for vehicle reuse (multiple flights)
and remote-site monitoring; (5) greater range, speed, and access to hazardous terrain than a surface rover; and (6) greater
resolution of surface details than an orbiter or balloons. Exploration of Titan presents an excellent opportunity for the
development and usage of such vehicles. Additional information is contained in the original extended abstract.
Derived from text
Vertical Takeoff Aircraft; Space Exploration; Powered Lift Aircraft; Rotary Wing Aircraft

20010041208 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA
Micro Navigator

Blaes, B. R.; Kia, T; Chau, S. N; Forum on Innovative Approaches to Outer Planetary Exploration 2001-2020; 2001, 9; In
English; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document

Miniature high-performance low-mass space avionics systems are desired for planned future outer planetary exploration
missions (i.e. Europa Orbiter/Lander, Pluto-Kuiper Express). The spacecraft fuel and mass requirements enabling orbit
insertion is the driving requirement. The Micro Navigator is an integrated autonomous Guidance, Navigation & Control
(GN&C)micro-system that would provide the critical avionics function for navigation, pointing, and precision landing. The
Micro Navigator hardware and software allow fusion of data from multiple sensors to provide a single integrated vehicle state
vector necessary for six degrees of freedom GN&C. The benefits of this MicroNavigator include: 1) The Micro Navigator
employs MEMS devices that promise orders of magnitude reductions in mass power and volume of inertial sensors
(accelerometers and gyroscopes), celestial sensing devices (startracker, sun sensor), and computing element; 2) The highly
integrated nature of the unit will reduce the cost of flight missions, a) The advanced miniaturization technologies employed

23

by the Micro Navigator lend themselves to mass production, and therefore will reduce production cost of spacecraft, b) The
integral approach simplifies interface issues associated with discrete components and reduces cost associated with integration
and test of multiple components; and 3) The integration of sensors and processing elements into a single unit will allow the
Micro Navigator to encapsulate attitude information and determination functions into a single object. This is particularly
beneficial for object-oriented software architectures that are used in advanced spacecraft. Additional information is contained
in the original extended abstract.
Derived from text
Autonomous Navigation; Miniaturization; Multisensor Fusion; Space Navigation; Spacecraft Guidance

20010038566 NASA Ames Research Center, Moffett Field, CA USA
International Agreement on Planetary Protection

Mars Sample Handling Protocol Workshop Series; October 2000, 93; In English; No Copyright; Avail: CASI; A01, Hardcopy
The maintenance of a NASA policy, is consistent with international agreements. The planetary protection policy
management in OSS, with Field Center support. The advice from internal and external advisory groups (NRC, NAC/Planetary
Protection Task Force). The technology research and standards development in bioload characterization. The technology
research and development in bioload reduction/sterilization. This presentation focuses on: forward contamination - research
on the potential for Earth life to exist on other bodies, improved strategies for planetary navigation and collision avoidance,
and improved procedures for sterile spacecraft assembly, cleaning and/or sterilization; and backward contamination -
development of sample transfer and container sealing technologies for Earth return, improvement in sample return landing
target assessment and navigation strategy, planning for sample hazard determination requirements and procedures, safety
certification, (liaison to NEO Program Office for compositional data on small bodies), facility planning for sample recovery
system, quarantine, and long-term curation of 4 returned samples.
Derived from text
International Cooperation; International Law; Policies; Planetary Environments; Environment Protection

20010024974 Colorado Univ., Boulder, CO USA
Low Velocity Impact Experiments in Microgravity

Colwell, J. E.; Sture, S.; Proceedings of the Fifth Microgravity Fluid Physics and Transport Phenomena Conference;
December 2000, 1335-1346; In English; No Copyright; Avail: CASI; A03, Hardcopy

Protoplanetary disks, planetary rings, the Kuiper belt, and the asteroid belt are collisionally evolved systems. Although
objects in each system may be bombarded by impactors at high interplanetary velocities, they are also subject to repeated
collisions at low velocities (v\hl0 m/s). In some regions of Saturn's rings, for example, the typical collision velocity inferred
from observations by the Voyager spacecraft and dynamical modeling is a fraction of a centimeter per second. These
interparticle collisions control the rate of energy dissipation in planetary rings and the rate of accretion in the early stages of
planetesimal formation. In the asteroid belt collisions typically occur at several km/s; however secondary craters are formed
at much lower impact speeds. In the Kuiper belt, where orbital speeds and eccentricities are much lower, collisions between
Kuiper belt objects (KBOs) can occur at speeds below 100 m/s. In the early solar system, KBOs accreted in the same way
planetesimals accreted in the inner solar system, however some regions of the Kuiper belt may now undergo erosional
collisions. Dust may be present on the surface of all of these objects in the form of a fine regolith created from micrometeoroid
bombardment (rings, asteroids, KBOs), high speed interparticle collisions (asteroids, KBOs) or as a product of accretion from
protoplanetary dust. Dust released in these collisions is often the only observable trace of the source objects and may be used
to infer the physical properties of those larger bodies. We are conducting a broad program of microgravity impact experiments
into dust to study the dissipation of energy in low energy collisions and the production of dust ejecta in these impacts. The
Collisions Into Dust Experiment (COLLIDE) flew on STS-90 in April 1998. The principal results of that experiment were
measurements of the coefficient of restitution for impacts into powders at impact speeds below 1 m/s. Almost no ejecta was
produced in impacts at 15 cm/s into JSC-1 powder, and the coefficient of restitution was about 0.03. COLLIDE-2 is undergoing
final preparations for a flight in 2001. The experiment will conduct six impact experiments at impact speeds between 1 and
100 cm/s. The target material will have a low relative density to mimic the regolith on low surface gravity objects in space,
such as planetesimals, planetary ring particles, and asteroids. A new experimental program, the Physics of Regolith Impacts
in Microgravity Experiment (PRIME) will use NASA's KC-135 aircraft to explore a much broader range of parameter space
than is possible with COLLIDE, at slightly higher impact velocities. PRIME will be capable of conducting up to 16 impact
experiments each flight day on the KC-135. Impact velocities between 50 cm/s and 5 m/s will be studied into a variety of target
materials and size distributions. The experiment will consist of an evacuated canister with 6 to 8 impact chambers on each
of two rotating turntables. Each impact chamber will include a target sample and a launcher with a unique set of parameters.

24

Two viewports will allow high speed video photography of impacts from two orthogonal views with the use of a mirror
mounted inside the canister. Data from COLLIDE and ground-based experimental studies suggest that particle size distribution
is an important parameter in controlling the response of granular media to low velocity impacts. Individual grain shapes may
also play an important role in the conversion of impactor kinetic energy to target grain kinetic energy. We will also make use
of numerical simulations of the impact process to understand the relevant parameters for experimental study. High speed video
of the impact and ejecta patterns will be analyzed to determine the ejecta mass and velocity distributions. This in turn will have
direct application for understanding the behavior of dust on the surfaces of planetary objects including asteroids and small
moons when disturbed by low velocity impacts and perturbations. These include naturally occurring impacts as well as
disturbances to the surface from human and spacecraft activity. The velocity distribution of the ejecta determines the amount
of material launched to various altitudes above the surface and escaping the parent body. This information is important for
spacecraft instruments landing on airless bodies with low surface gravity and powdery regoliths.
Author (revised)
Collisions; Dust; Gravitational Effects; Low Speed; Microgravity; Impact; Energy Dissipation

20010023145 Centre National d'Etudes Spatiales, France
The Stakes of the Aerocapture for Missions to Mars

Cledassou, R.; Lam-Trong, Th.; Charbonnier, J. M.; Concepts and Approaches for Mars Exploration; July 2000, Issue Part 1,
186; In English; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document

The Hohmann transfer trajectory is an economical way to go from Earth to Mars but a spacecraft has to reduce its speed
very significantly upon arrival in order to be inserted into a Mars orbit. The aerocapture is a way to do that, by using the
Martian atmosphere to produce sufficient aerodynamic drag force on a heatshield and achieve the required deceleration. This
presentation will address the major stake of the aerocapture which is twofold: a) We will list the different technologies and
areas of knowledge related to the aerocapture, identify the risks associated with each of them and finally demonstrate that
aerocapture is not as risky as it is said to be; b) Aerocapture saves a huge amount of propellant and so allows to improve
dramatically the dollar/kg ratio for any payload at Mars by using this mass savings for payloads and by decreasing the launch
cost. This benefit is particularly evident for a return mission because of the amplification factor of the propellant mass for the
escape of Mars ('snow ball' effect). We will have a quantitative analysis of some typical cases of spacecraft vs. launcher
performance . We will conclude that aerocapture is interesting for the present robotic missions and certainly a good investment
for the future manned missions to Mars.
Derived from text
Mars Missions; Aerocapture; Earth-Mars Trajectories; Transfer Orbits

20010023141 Science Applications International Corp., Littleton, CO USA
Precision Terminal Guidance for a Mars Lander

Klarquist, William N.; Wahl, Beth E.; Lowrie, James W.; Concepts and Approaches for Mars Exploration; July 2000, Issue
Part 1, 178-179; In English; No Copyright; Avail: CASI; A01, Hardcopy

To date Mars landers have relied solely on Earth-based navigation measurements to achieve a desired landing site.
They've had no active guidance and control system to monitor and control the entry and descent trajectory or guide the final
landing. This results in very large landing site uncertainties (\gl80 km x 20 km) and precludes targeting specific, small scale
regions such as canyons and flood channels. Moreover, localized hazards cannot be sensed or avoided, resulting in higher
mission risk. SAIC's Center for Intelligent Systems, (SAIC-CIS) based on current and past research, believes that reliably
accurate landings at pre-selected sites are achievable and that the mission risk associated with local hazards can be greatly
reduced. Our concept involves applying an integrated system level solution that leverages the tremendous amount of
information available on the Martian environment and applies modern technologies in the areas of visual based navigation,
maneuverable parachutes, and advanced sensors.
Derived from text
Active Control; Terminal Guidance; Spacecraft Guidance; Mars Landing

20010023136 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA
Mars Sample Return without Landing on the Surface

Jurewicz, A. J. G.; Jones, Steven M.; Yen, A. S.; Concepts and Approaches for Mars Exploration; July 2000, Issue Part 1,
168-169; In English; No Copyright; Avail: CASI; A01, Hardcopy

Many in the science community want a Mars sample return in the near future, with the expectation that it will provide

25

in-depth information, significantly beyond what we know from remote sensing, limited in-situ measurements, and work with
Martian meteorites. Certainly, return of samples from the Moon resulted in major advances in our understanding of both the
geologic history of our planetary satellite, and its relationship to Earth. Similar scientific insights would be expected from
analyses of samples returned from Mars. Unfortunately, Mars-lander sample-return missions have been delayed, for the reason
that NASA needs more time to review the complexities and risks associated with that type of mission. A traditional sample
return entails a complex transfer-chain, including landing, collection, launch, rendezvous, and the return to Earth, as well as
an evaluation of potential biological hazards involved with bringing pristine Martian organics to Earth. There are, however,
means of returning scientifically-rich samples from Mars without landing on the surface. This paper discusses an approach for
returning intact samples of surface dust, based on known instrument technology, without using an actual Martian lander.
Derived from text
Mars Sample Return Missions; Particulate Sampling; Aerocapture

20010020516 Arizona Univ., Tucson, AZ USA
The Martian Oasis Detector

Smith, P. H.; tomasko, M. G.; McEwen, A.; Rice, J.; Concepts and Approaches for Mars Exploration; July 2000, Issue Part
2, 286-287; In English; No Copyright; Avail: CASI; A01, Hardcopy

The next phase of unmanned Mars missions paves the way for astronauts to land on the surface of Mars. There are lessons
to be learned from the unmanned precursor missions to the Moon and the Apollo lunar surface expeditions. These unmanned
missions (Ranger, Lunar Orbiter, and Surveyor) provided the following valuable information, useful from both a scientific and
engineering perspective, which was required to prepare the way for the manned exploration of the lunar surface: (1) high
resolution imagery instrumental to Apollo landing site selection also tremendously advanced the state of Nearside and Farside
regional geology; (2) demonstrated precision landing (less than two kilometers from target) and soft landing capability; (3)
established that the surface had sufficient bearing strength to support a spacecraft; and (4) examination of the chemical
composition and mechanical properties of the surface. The search for extinct or extant life on Mars will follow the water.
However, geomorphic studies have shown that Mars has had liquid water on its surface throughout its geologic history. A
cornucopia of potential landing sites with water histories (lakes, floodplains, oceans, deltas, hydrothermal regions) presently
exist. How will we narrow down site selection and increase the likelihood of finding the signs of life? One way to do this is
to identify 'Martian oases.' It is known that the Martian surface is often highly fractured and some areas have karst structures
that support underground caves. Much of the water that formed the channels and valley networks is thought to be frozen
underground. All that is needed to create the potential for liquid water is a near surface source of heat; recent lava flows and
Martian meteorites attest to the potential for volcanic activity. If we can locate even one spot where fracturing, ice, and
underground heat are co-located then we have the potential for an oasis. Such a discovery could truly excite the imaginations
of both the public and Congress providing an attainable goal for both robotic and manned missions. The instrument required
to detect an active oasis is a high spatial resolution (few tens of meters) Short Wavelength Infrared (SWIR) spectrometer
coupled with a high resolution camera (five m/pixel). This combination creates too large a data volume to possibly return data
for the entire Martian Surface; therefore it has been designed as one of the first in a new generation of 'smart' detectors, called
the Mars Oasis Detector (MOD).
Author
Mars (Planet); Water; Mars Missions; Mars Surface; Spacecraft Instruments; Oases; Mars Exploration

20010020495 California Univ., Los Angeles, CA USA
After the Mars Polar Lander: Where to Next?

Paige, D. A.; Boynton, W. V.; Crisp, D.; DeJong, E.; Hansen, C. J.; Harri, A. M.; Keller, H. U.; Leshin, L. A.; May, R. D.;
Smith, P. H., et al.; Concepts and Approaches for Mars Exploration; July 2000, Issue Part 2, 245-246; In English; No
Copyright; Avail: CASI; A01, Hardcopy

The recent loss of the Mars Polar Lander (MPL) mission represents a serious setback to Mars science and exploration.
Targeted to land on the Martian south polar layered deposits at 76 degrees south latitude and 195 degrees west longitude, it
would have been the first mission to study the geology, atmospheric environment, and volatiles at a high-latitude landing site.
Since the conception of the MPL mission, a Mars exploration strategy has emerged which focuses on Climate, Resources and
Life, with the behavior and history of water as the unifying theme. A successful MPL mission would have made significant
contributions towards these goals, particularly in understanding the distribution and behavior of near-surface water, and the
nature and climate history of the south polar layered deposits. Unfortunately, due to concerns regarding the design of the MPL
spacecraft, the rarity of direct trajectories that enable high-latitude landings, and funding, an exact reflight of MPL is not
feasible within the present planning horizon. However, there remains significant interest in recapturing the scientific goals of

26

the MPL mission. The following is a discussion of scientific and strategic issues relevant to planning the next polar lander

mission, and beyond.

Author

Mars Exploration; Mission Planning; Mars Missions; Polar Regions; Mars (Planet)

20010020467 Science Applications International Corp., Littleton, CO USA
Precision Navigation for a Mars Airplane

Lowrie, James W.; Concepts and Approaches for Mars Exploration; July 2000, Issue Part 2, 196-197; In English; No
Copyright; Avail: CASI; A01, Hardcopy

The rough Martian terrain significantly impedes high speed travel by wheeled vehicles and much of it is simply
inaccessible given the capability of typical rover designs. Airplanes, however, have much greater range and can provide access
to scientifically interesting terrain that is inaccessible to landers and rovers. Moreover, they can provide coverage of a large
portion of the surface and return high resolution images and science data not practical from orbiting spacecraft. Precise
navigation on Earth requires a constellation of satellites such as GPS (Global Positioning Satellites) or a network of precisely
located and calibrated ground beacons, an approach that is impractical for Mars exploration in the near future. In order to
realize the benefits of airplane exploration on Mars, a precision navigation system is required. Such a system also provides
a high degree of autonomous capability because it enables: (1) Accurate overflight of specifically targeted sites. (2) Hazard
avoidance in low altitude flight. (3) The collection of 'focused' science data which reduces overall data volume and supports
an optimized data return strategy (4) Accurate spatial and temporal correlation of acquired science data with orbiter
observations. (5) A geodetically referenced site survey capability. (6) A soft landing capability by providing in-flight landing
site selection and terminal guidance. (7) Return to a base station following flight. (8) Precise placement of science probes and
future navigation beacons. SAIC's Center for Intelligent Systems (SAIC-CIS) leverages on experience from unmanned vehicle
research to propose a concept for an intelligent landmark navigation system that relies on autonomous real-time recognition
of visible surface features during flight.
Author
Mars Missions; Mars Exploration; Mars (Planet); Autonomous Navigation; Aircraft

20010019289 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA
MOLA-Based Landing Site Characterization

Duxbury, T. C; Ivanov, A. B.; First Landing Site Workshop for the 2003 Mars Exploration Rovers; 2001, 18-19; In English;
No Copyright; Abstract Only; Available from CASI only as part of the entire parent document

The Mars Global Surveyor (MGS) Mars Orbiter Laser Altimeter (MOLA) data provide the basis for site characterization
and selection never before possible. The basic MOLA information includes absolute radii, elevation and 1 micrometer albedo
with derived datasets including digital image models (DIM's illuminated elevation data), slopes maps and slope statistics and
small scale surface roughness maps and statistics. These quantities are useful in downsizing potential sites from descent
engineering constraints and landing/roving hazard and mobility assessments. Slope baselines at the few hundred meter level
and surface roughness at the 10 meter level are possible. Additionally, the MOLA-derived Mars surface offers the possibility
to precisely register and map project other instrument datasets (images, ultraviolet, infrared, radar, etc.) taken at different
resolution, viewing and lighting geometry, building multiple layers of an information cube for site characterization and
selection. Examples of direct MOLA data, data derived from MOLA and other instruments data registered to MOLA arc given
for the Hematite area.
Author
Mars Global Surveyor; Mars Surface; Landing Sites; Laser Altimeters

20010002491 California Univ., Los Angeles, CA USA
After the Mars Polar Lander: Where to Next?

Paige, D. A.; International Conference on Mars Polar Science and Exploration; August 2000, 140-141; In English; No
Copyright; Avail: CASI; A01, Hardcopy

The recent loss of the Mars Polar Lander (MPL) mission represents a serious setback to Mars science and exploration.
Targeted to land on the Martian south polar layered deposits at 76 deg south latitude and 195 deg west longitude, it would
have been the first mission to study the geology, atmospheric environment, and volatiles at a high-latitude landing site. Since
the conception of the MPL mission, a Mars exploration strategy has emerged which focuses on Climate, Resources and Life,
with the behavior and history of water as the unifying theme. A successful MPL mission would have made significant

27

contributions towards these goals, particularly in understanding the distribution and behavior of near-surface water, and the
nature and climate history of the south polar layered deposits. Unfortunately, due to concerns regarding the design of the MPL
spacecraft, the rarity of direct trajectories that enable high-latitude landings, and funding, an exact reflight of MPL is not
feasible within the present planning horizon. However, there remains significant interest in recapturing the scientific goals of
the MPL mission. The following is a discussion of scientific and strategic issues relevant to planning the next polar lander
mission, and beyond. Additional information is contained in the original extended abstract.
Author
Mars (Planet); Mars Exploration; Mars Polar Lander; Polar Regions; Mars Missions

20000102372 Tennessee Univ., Knoxville, TN USA

Earth Return Aerocapture for the TransHab/Ellipsled Vehicle

Muth, W. D.; Hoffmann, G; Lyne, J. E.; October 2000; In English
Contract(s)/Grant(s): NAG 1-2 163; No Copyright; Avail: CASI; A04, Hardcopy

The current architecture being considered by NASA for a human Mars mission involves the use of an aerocapture
procedure at Mars arrival and possibly upon Earth return. This technique would be used to decelerate the vehicles and insert
them into their desired target orbits, thereby eliminating the need for propulsive orbital insertions. The crew may make the
interplanetary journey in a large, inflatable habitat known as the TransHab. It has been proposed that upon Earth return, this
habitat be captured into orbit for use on subsequent missions. In this case, the TransHab would be complimented with an
aeroshell, which would protect it from heating during the atmospheric entry and provide the vehicle with aerodynamic lift.
The aeroshell has been dubbed the 'Ellipsled' because of its characteristic shape. This paper reports the results of a preliminary
study of the aerocapture of the TransHab/Ellipsled vehicle upon Earth return. Undershoot and overshoot boundaries have been
determined for a range of entry velocities, and the effects of variations in the atmospheric density profile, the vehicle
deceleration limit, the maximum vehicle roll rate, the target orbit, and the vehicle ballistic coefficient have been examined. A
simple, 1 80 degree roll maneuver was implemented in the undershoot trajectories to target the desired 407 km circular Earth
orbit. A three-roll sequence was developed to target not only a specific orbital energy, but also a particular inclination, thereby
decreasing propulsive inclination changes and post-aerocapture delta-V requirements. Results show that the TransHab/
Ellipsled vehicle has a nominal corridor width of at least 0.7 degrees for entry speeds up to 14.0 km/s. Most trajectories were
simulated using continuum flow aerodynamics, but the impact of high-altitude viscous effects was evaluated and found to be
minimal. In addition, entry corridor comparisons have been made between the TransHab/Ellipsled and a modified Apollo
capsule which is also being considered as the crew return vehicle; because of its slightly higher lift-to-drag ratio, the TransHab
has a modest advantage with regard to corridor width. Stagnation-point heating rates and integrated heat loads were
determined for a range of vehicle ballistic coefficients and entry velocities.
Author

Spacecraft Design; Product Development; Aerocapture; Aeromaneuvering; Interplanetary Transfer Orbits; Atmospheric
Entry

20000085950 North Carolina State Univ., Raleigh, NC USA

An Investigation of Terminal Guidance and Control Techniques for a Robotic Mars Lander

Birge, Brian K.; Walberg, Gerald; [2000]; In English

Contract(s)/Grant(s): NAG1-2222; No Copyright; Avail: CASI; A03, Hardcopy

Continuing on previous work, various precision landing control algorithms arc examined with the goal of minimizing the
landed distance to a specified location on the Mars surface. This study considers a set of points from parachute handoff to
touchdown on the surface. The first scenario considers a reverse gravity turn to a hover condition 500 meters above the surface
and then uses lateral thrusting to minimize die range to target. The second scenario examines a guided, lifting parachute
followed by a powered gravity turn to the targeted landing site. The third scenario considers thrust vectoring while on the
ballistic parachute, followed by a reverse gravity turn to touchdown.
Author
Terminal Guidance; Command Guidance; Control Systems Design; Thrust Vector Control; Control Theory

20000074639 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA

The Deep Space 4/Champollion Comet Rendezvous and Lander Technology Demonstration Mission

Smythe, William D.; Weissman, Paul R.; Muirhead, Brian K.; Tan-Wang, Grace H.; Sabahi, Dara; Grimes, James M.; [2000];
In English; No Copyright; Avail: Other Sources; Abstract Only

28

The Deep Space 4/Champollion mission is designed to test and validate technologies for landing on and anchoring to
small bodies, and sample collection and transfer, in preparation for future sample return missions from comets, asteroids, and
satellites, in addition, DS-4 will test technologies for advanced, multi-engine solar electric propulsion (SEP) systems,
inflatable-rigidizable solar arrays, autonomous navigation and precision guidance for landing, autonomous hazard detection
and avoidance, and advanced integrated avionics and packaging concepts. Deep Space-4/Champollion consists of two
spacecraft: an orbiter/carrier vehicle which includes the multi-engine SEP stage, and a lander, called Champollion, which will
descend to the surface of the 46P/Tempel 1 cometary nucleus. The spacecraft will launch in April, 2003 and land on the comet
in September, 2006 Deep Space 4/Champollion is a joint project between NASA and CNES, the French space agency.
Author
Deep Space; Space Missions; Mission Planning; Landing; Comets; Asteroids

20000074247 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA
Navigation Strategy for the Mars 2001 Lander Mission

Mase, Robert A.; Spencer, David A.; Smith, John C; Braun, Robert D.; [2000]; In English; No Copyright; Avail: Other
Sources; Abstract Only

The Mars Surveyor Program (MSP) is an ongoing series of missions designed to robotically study, map and search for
signs of life on the planet Mars. The MSP 2001 project will advance the effort by sending an orbiter, a lander and a rover to
the red planet in the 2001 opportunity. Each vehicle will carry a science payload that will Investigate the Martian environment
on both a global and on a local scale. Although this mission will not directly search for signs of life, or cache samples to be
returned to Earth, it will demonstrate certain enabling technologies that will be utilized by the future Mars Sample Return
missions. One technology that is needed for the Sample Return mission is the capability to place a vehicle on the surface
within several kilometers of the targeted landing site. The MSP'01 Lander will take the first major step towards this type of
precision landing at Mars. Significant reduction of the landed footprint will be achieved through two technology advances. The
first, and most dramatic, is hypersonic aeromaneuvering; the second is improved approach navigation. As a result, the guided
entry will produce in a footprint that is only tens of kilometers, which is an order of magnitude improvement over the
Pathfinder and Mars Polar Lander ballistic entries. This reduction will significantly enhance scientific return by enabling the
potential selection of otherwise unreachable landing sites with unique geologic interest and public appeal. A landed footprint
reduction from hundreds to tens of kilometers is also a milestone on the path towards human exploration of Mars, where the
desire is to place multiple vehicles within several hundred meters of the planned landing site. Hypersonic aeromaneuvering
is an extension of the atmospheric flight goals of the previous landed missions, Pathfinder and Mars Polar Lander (MPL), that
utilizes aerodynamic lift and an autonomous guidance algorithm while in the upper atmosphere. The onboard guidance
algorithm will control the direction of the lift vector, via bank angle modulation, to keep the vehicle on the desired trajectory.
While numerous autonomous guidance algorithms have been developed for use during hypersonic flight at Earth, this will be
the first flight of an autonomously directed lifting entry vehicle at Mars. However, without sufficient control and knowledge
of the atmospheric entry conditions, the guidance algorithm will not perform effectively. The goal of the interplanetary
navigation strategy is to deliver the spacecraft to the desired entry condition with sufficient accuracy and knowledge to enable
satisfactory guidance algorithm performance. Specifically, the entry flight path angle must not exceed 0.27 deg. to a 3 sigma
confidence level. Entry errors will contribute directly to the size of the landed footprint and the most significant component
is entry flight path angle. The size of the entry corridor is limited on the shallow side by integrated heating constraints, and
on the steep side by deceleration (g-load) and terminal descent propellant. In order to meet this tight constraint it is necessary
to place a targeting maneuver seven hours prior to the time of entry. At this time the trajectory knowledge will be quite
accurate, and the effects of maneuver execution errors will be small. The drawback is that entry accuracy is dependent on the
success of this final late maneuver. Because propulsive maneuvers are critical events, it is desirable to minimize their
occurrence and provide the flight team with as much response time as possible in the event of a spacecraft fault. A mission
critical maneuver at Entry - 7 hours does not provide much fault tolerance, and it is desirable to provide a strategy that
minimizes reliance on this maneuver. This paper will focus on the Improvements in interplanetary navigation that will decrease
entry errors and will reduce the landed footprint, even in the absence of aeromaneuvering. The easiest to take advantage of
are Improvements In the knowledge of the Mars ephemeris and gravity field due to the MGS and MSP'98 missions.
Improvements In data collection and reduction techniques such as 'precision ranging' and near-simultaneous tracking will also
be utilized. In addition to precise trajectory control, a robust strategy for communications and flight operations must also be
demonstrated. The result Is a navigation and communications strategy on approach that utilizes optimal maneuver placement
to take advantage of trajectory knowledge, minimizes risk for the flight operations team, is responsive to spacecraft hardware
limitations, and achieves the entry corridor. The MSP2001 mission Is managed at JPL under the auspices of the Mars

29

Exploration Directorate. The spacecraft flight elements are built and managed by Lockheed-Martin Astronautics in Denver,

Colorado.

Author

Interplanetary Navigation; Landing Sites; Mars Landing; Earth-Mars Trajectories; Orbital Mechanics; Orbit Calculation;

Mars Surveyor 2001 Mission

20000074083 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA
A Light- Weight Inflatable Hypersonic Drag Device for Planetary Entry

McRonald, Angus D.; [2000]; In English; No Copyright; Avail: Other Sources; Abstract Only

The author has analyzed the use of a light-weight inflatable hypersonic drag device, called a ballute, for flight in planetary
atmospheres, for entry, aerocapture, and aerobraking. Studies to date include Mars, Venus, Earth, Saturn, Titan, Neptune and
Pluto, and data on a Pluto lander and a Mars orbiter will be presented to illustrate the concept. The main advantage of using
a ballute is that aero, deceleration and heating in atmospheric entry occurs at much smaller atmospheric density with a ballute
than without it. For example, if a ballute has a diameter 10 times as large as the spacecraft, for unchanged total mass, entry
speed and entry angle,the atmospheric density at peak convective heating is reduced by a factor of 100, reducing the heating
by a factor of 10 for the spacecraft and a factor of 30 for the ballute. Consequently the entry pay load (lander, orbiter, etc) is
subject to much less heating, requires a much reduced thermal, protection system (possibly only an MLI blanket), and the
spacecraft design is therefore relatively unchanged from its vacuum counterpart. The heat flux on the ballute is small enough
to be radiated at temperatures below 800 K or so. Also, the heating may be reduced further because the ballute enters at a more
shallow angle, even allowing for the increased delivery angle error. Added advantages are less mass ratio of entry system to
total entry mass, and freedom from the low-density and transonic instability problems that conventional rigid entry bodies
suffer, since the vehicle attitude is determined by the ballute, usually released at continuum conditions (hypersonic for an
orbiter, and subsonic for a lander). Also, for a lander the range from entry to touchdown is less, offering a smaller footprint.
The ballute derives an entry corridor for aerocapture by entering on a path that would lead to landing, and releasing the ballute
adaptively, responding to measured deceleration, at a speed computed to achieve the desired orbiter exit conditions. For a
lander an accurate landing point could be achieved by providing the lander with a small gliding capacity, using the large
potential energy available from being subsonic at high altitude. Alternatively the ballute can be retained to act as a parachute
or soft-landing device, or to float the payload as a buoyant aerobot. As expected, the ballute has smaller size for relatively
small entry speeds, such as for Mars and Titan, or for the extensive atmosphere of a low-gravity planet such as Pluto. Details
of a ballute to place a small Mars orbiter and a small Pluto lander will be given to illustrate the concept. The author will discuss
presently available ballute materials and a development program of aerodynamic tests and materials that would be required
for ballutes to achieve their full potential.
Author

Aerodynamic Heating; Research; Ballutes; Buoyancy; Drag Devices; Floats; Inflatable Structures; Microgravity; Planetary
Atmospheres; Spacecraft Design

20000062309 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA
Precise Image-Based Motion Estimation for Autonomous Small Body Exploration

Johnson, Andrew Edie; Matthies, Larry H.; [2000]; In English; 5th; No Copyright; Avail: CASI; A01, Hardcopy

We have developed and tested a software algorithm that enables onboard autonomous motion estimation near small bodies
using descent camera imagery and laser altimetry. Through simulation and testing, we have shown that visual feature tracking
can decrease uncertainty in spacecraft motion to a level that makes landing on small, irregularly shaped, bodies feasible.
Possible future work will include qualification of the algorithm as a flight experiment for the Deep Space 4/Champollion comet
lander mission currently under study at the Jet Propulsion Laboratory.
Author
Estimating; Autonomy; Spacecraft Motion; Optical Tracking; Image Analysis; Algorithms

20000057306 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA

Aerobraking at Venus and Mars: A Comparison of the Magellan and Mars Global Surveyor Aerobraking Phases

Lyons, Daniel T; [2000]; In English; No Copyright; Avail: CASI; A01, Hardcopy

On February 4, 1999 the Mars Global Surveyor spacecraft became the second spacecraft to successfully aerobrake into
a nearly circular orbit about another planet. This paper will highlight some of the similarities and differences between the
aerobraking phases of this mission and the first mission to use aerobraking, the Magellan mission to Venus. Although the Mars

30

Global Surveyor (MGS) spacecraft was designed for aerobraking and the Magellan spacecraft was not, aerobraking MGS was
a much more challenging task than aerobraking Magellan, primarily because the spacecraft was damaged during the initial
deployment of the solar panels. The MGS aerobraking phase had to be completely redesigned to minimize the bending
moment acting on a broken yoke connecting one of the solar panels to the spacecraft. Even if the MGS spacecraft was
undamaged, aerobraking at Mars was more challenging than aerobraking at Venus for several reasons. First, Mars is subject
to dust storms, which can significantly change the temperature of the atmosphere due to increased solar heating in the low and
middle altitudes (below 50 km), which in turn can significantly increase the density at the aerobraking altitudes (above 100
km). During the first part of the MGS aerobraking phase, a regional dust storm was observed to have a significant and very
rapid effect on the entire atmosphere of Mars. Computer simulations of global dust storms on Mars indicate that even larger
density increases are possible than those observed during the MGS aerobraking phases. For many aerobraking missions, the
duration of the aerobraking phase must be kept as short as possible to minimize the total mission cost. For Mars missions, a
short aerobraking phase means that there will be less margin to accommodate atmospheric variability, so the operations team
must be ready to propulsively raise periapsis by tens of kilometers on very short notice. This issue was less of a concern on
Venus, where the thick lower atmosphere and the slow planet rotation resulted in more predictable atmospheric densities from
one orbit to the next.
Author
Aerobraking; Mars Global Surveyor; Magellan Spacecraft (NASA); Circular Orbits

20000056881 Jet Propulsion Lab., California Inst, of Tech., Pasadena, CA USA
The Strategy for the Second Phase of Aerobraking Mars Global Surveyor

Johnston, M. D.; Esposito, P. B.; Alwar, V; Demcak, S. W.; Graat, E. J.; Burkhart, P. D.; Portock, B. M.; [2000]; In English;
No Copyright; Avail: CASI; A01, Hardcopy

On February 19, 1999, the Mars Global Surveyor (MGS) spacecraft was able to propulsively establish its mapping orbit.
This event followed the completion of the second phase of aerobraking for the MGS spacecraft on February 4, 1999. For the
first time, a spacecraft at Mars had successfully employed aerobraking methods in order to reach its desired pre-launch
mapping orbit. This was accomplished despite a damaged spacecraft solar array. The MGS spacecraft was launched on
November 7, 1996, and after a ten month interplanetary transit was inserted into a highly elliptical capture orbit at Mars on
September 12, 1997. Unlike other interplanetary missions, the MGS spacecraft was launched with a planned mission delta-V
((Delta) V) deficit of nearly 1250 m/s. To overcome this AV deficit, aerobraking techniques were employed. However, damage
discovered to one of the spacecraft's two solar arrays after launch forced major revisions to the original aerobraking planning
of the MGS mission. In order to avoid a complete structural failure of the array, peak dynamic pressure levels for the spacecraft
were established at a major spacecraft health review in November 1997. These peak dynamic pressure levels were roughly
one-third of the original mission design values. Incorporating the new dynamic pressure limitations into mission replanning
efforts resulted in an 'extended' orbit insertion phase for the mission. This 'extended' orbit insertion phase was characterized
by two distinct periods of aerobraking separated by an aerobraking hiatus that would last for several months in an intermediate
orbit called the 'Science Phasing Orbit' (SPO). This paper describes and focuses on the strategy for the second phase of
aerobraking for the MGS mission called 'Aerobraking Phase 2.' This description will include the baseline aerobraking flight
profile, the trajectory control methodology, as well as the key trajectory metrics that were monitored in order to successfully
'guide' the spacecraft to its desired mapping orbit. Additionally, the actual aerobraking progress is contrasted to the planned
aerobraking flight profile. (A separate paper will describe the navigation aspects of MGS aerobraking in detail.) Key to the
success of the MGS mission is the delivery of the spacecraft to its final mapping orbit and the synergy the instrument
complement provides to its scientific investigators when science data is returned from that orbit. The MGS mapping orbit is
characterized as a low altitude, near-circular, near-polar orbit that is Sun-synchronous with the descending equatorial crossing
at 2:00 AM local mean solar time (LMST).
Derived from text
Aerobraking; Mars Global Surveyor; Mapping; Orbit Insertion; Trajectory Control; Navigation; Mission Planning

31

Subject Terms

2001 MARS ODYSSEY

Application of Accel era meter Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

Plume Modeling and Application to Mars
2001 Odyssey Aerobraking - 16

Thermal Analysis and Correlation of the
Mars Odyssey Spacecraft's Solar Array
During Aerobraking Operations - 17

ACCELEROMETERS

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

Approaches to autonomous aerobraking
at Mars - 1 5

ACOUSTO-OPTICS

AIMS: Acousto-optic imaging spectrom-
eter for spectral mapping of solid
surfaces - 13

ACTIVE CONTROL

Precision Terminal Guidance for a Mars
Lander - 25

AEROASSIST

Aeroassist Technology Planning for
Exploration - 4

Aerothermal Instrumentation Loads To
Implement Aeroassist Technology in Fu-
ture Robotic and Human Missions to
MARS and Other Locations Within the
Solar System - 20

Atmospheric Models for Aeroentry and
Aeroassist - 1

NASA Development of Aerocapture
Technologies - 14

Study of Orbital Transfers with Aeroas-
sisted Maneuvers - 21

AEROBRAKING

Aerobraking at Venus and Mars: A Com-
parison of the Magellan and Mars Global
Surveyor Aerobraking Phases - 30

Aeroheating Thermal Analysis Methods
for Aerobraking Mars Missions - 4

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

Atmospheric Models for Aerocapture
Systems Studies - 1

Atmospheric Models for Aeroentry and
Aeroassist - 1

Autonomous Aerobraking at Mars - 16

Plume Modeling and Application to Mars
2001 Odyssey Aerobraking - 16

The Development and Evaluation of an
Operational Aerobraking Strategy for the
Mars 2001 Odyssey Orbiter - 17

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

AEROCAPTURE

Aerocapture Guidance Algorithm Com-
parison Campaign - 18

Aerocapture Performance Analysis for a
Neptune-Triton Exploration Mission - 2

Aerocapture Technology Development
Needs for Outer Planet Exploration - 20

Aerocapture Technology Project
Overview - 11

Angle-of-Attack-Modulated Terminal
Point Control for Neptune Aerocapture

- 9

Atmospheric Models for Aerocapture
Systems Studies - 1

Atmospheric Models for Aerocapture - 4

Atmospheric Models for Aeroentry and
Aeroassist - 1

Connecting Atmospheric Science and At-
mospheric Models for Aerocaptured Mis-
sions to Titan and the Outer Planets - 5

Earth Return Aerocapture for the
TransHab/Ellipsled Vehicle - 28

Engineering-Level Model Atmospheres
for Titan & Neptune - 13

Mars Sample Return without Landing on
the Surface - 25

NASA Development of Aerocapture
Technologies - 14

Neptune Aerocapture Systems Analysis

- 3

Preliminary Convective-Radiative Heat-
ing Environments for a Neptune Aero-
capture Mission - 1

Radioisotope Electric Propulsion for Fast
Outer Planetary Orbiters - 20

Structural Design for a Neptune Aerocap-
ture Mission - 2

The Stakes of the Aerocapture for Mis-
sions to Mars - 25

Trailing Ballute Aerocapture: Concept
and Feasibility Assessment - 12

Wake Closure Characteristics and After-
body Heating on a Mars Sample Return
Orbiter- 15

AERODYNAMIC CHARACTERISTICS

Experimental Hypersonic Aerodynamic
Characteristics of the 2001 Mars Sur-
veyor Precision Lander with Flap - 18

Study of Orbital Transfers with Aeroas-
sisted Maneuvers - 21

AERODYNAMIC DRAG

Trailing Ballute Aerocapture: Concept
and Feasibility Assessment - 12

AERODYNAMIC FORCES

NASA Development of Aerocapture
Technologies - 14

AERODYNAMIC HEATING

A Light-Weight Inflatable Hypersonic
Drag Device for Planetary Entry - 30

Aeroheating Thermal Analysis Methods
for Aerobraking Mars Missions - 4

Control Surface and Afterbody Experi-
mental Aeroheating for a Proposed Mars
Smart Lander Aeroshell - 17

Thermal Analysis and Correlation of the
Mars Odyssey Spacecraft's Solar Array
During Aerobraking Operations - 17

Wake Closure Characteristics and After-
body Heating on a Mars Sample Return
Orbiter- 15

AERODYNAMIC LOADS

Aerothermal Instrumentation Loads To
Implement Aeroassist Technology in Fu-
ture Robotic and Human Missions to
MARS and Other Locations Within the
Solar System - 20

AERODYNAMICS

Pitch control during autonomous aero-
braking for near-term Mars exploration
- 12

AEROMANEUVERING

Autonomous Aerobraking at Mars

16

Earth Return Aerocapture for the
TransHab/Ellipsled Vehicle - 28

AEROSHELLS

Control Surface and Afterbody Experi-
mental Aeroheating for a Proposed Mars
Smart Lander Aeroshell - 17

Neptune Aerocapture Systems Analysis

- 3

AEROSPACE SCIENCES

AIMS: Acousto-optic imaging spectrom-
eter for spectral mapping of solid
surfaces - 13

Blended control, predictor-corrector guid-
ance algorithm: An enabling technology
for Mars aerocapture - 10

Daily repeat-groundtrack Mars orbits - 7

Entry descent, and landing scenario for
the Mars exploration Rover mission - 7

Europa Lander - 13

Pioneer Venus and Galileo entry probe
heritage - 8

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system

- 10

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Thermal protection system technology
and facility needs for demanding future
planetary missions - 8

A-1

AEROTHERMODYNAMICS

Aerocapture Technology Development
Needs for Outer Planet Exploration - 20

Aerothermal Instrumentation Loads To
Implement Aeroassist Technology in Fu-
ture Robotic and Human Missions to
MARS and Other Locations Within the
Solar System - 20

Experimental Hypersonic Aerodynamic
Characteristics of the 2001 Mars Sur-
veyor Precision Lander with Flap - 18

Preliminary Convective-Radiative Heat-
ing Environments for a Neptune Aero-
capture Mission - 1

AIR BAG RESTRAINT DEVICES

Entry trajectory and atmosphere recon-
struction methodologies for the mars ex-
ploration rover mission - 7

AIRCRAFT

Precision Navigation for a Mars Airplane

- 27

ALGORITHMS

Aerocapture Guidance Algorithm Com-
parison Campaign - 18

Aerocapture Guidance Methods for High
Energy Trajectories - 11

Angle-of-Attack-Modulated Terminal
Point Control for Neptune Aerocapture

- 9

Precise Image-Based Motion Estimation
for Autonomous Small Body Exploration

- 30

ANGLE OF ATTACK

Angle-of-Attack-Modulated Terminal
Point Control for Neptune Aerocapture

- 9

ANNUAL VARIATIONS

Connecting Atmospheric Science and At-
mospheric Models for Aerocaptured Mis-
sions to Titan and the Outer Planets - 5

APPLICATIONS PROGRAMS (COMPUT-
ERS)

Study of Orbital Transfers with Aeroas-
sisted Maneuvers - 21

ASTEROIDS

The Deep Space 4/Champollion Comet
Rendezvous and Lander Technology
Demonstration Mission - 28

ASTROPHYSICS

Blended control, predictor-corrector guid-
ance algorithm: An enabling technology
for Mars aerocapture - 10

ATMOSPHERIC CHEMISTRY

Blended control, predictor-corrector guid-
ance algorithm: An enabling technology
for Mars aerocapture - 10

ATMOSPHERIC DENSITY

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

Connecting Atmospheric Science and At-
mospheric Models for Aerocaptured Mis-
sions to Titan and the Outer Planets - 5

Development of a Monte Carlo Mars-
gram model for 2001 Mars Odyssey
aerobraking simulations - 15

ATMOSPHERIC ENTRY

CFD Prediction of the BEAGLE 2 Mars
Probe Aerodynamic Database - 19

Computational Analysis of Towed Ballute
Interactions - 18

Earth Return Aerocapture for the
TransHab/Ellipsled Vehicle - 28

Mars Exploration Rover Six-Degree-Of-
Freedom Entry Trajectory Analysis - 1 4

Multibody Parachute Flight Simulations
for Planetary Entry Trajectories Using
'Equilibrium Points' - 16

Uncertainty Optimization Applied to the
Monte Carlo Analysis of Planetary Entry
Trajectories - 23

ATMOSPHERIC MODELS

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

Atmospheric Models for Aerocapture
Systems Studies - 1

Atmospheric Models for Aerocapture - 4

Atmospheric Models for Aeroentry and
Aeroassist - 1

Connecting Atmospheric Science and At-
mospheric Models for Aerocaptured Mis-
sions to Titan and the Outer Planets - 5

Engineering-Level Model Atmospheres
for Titan & Neptune - 13

ATMOSPHERIC PHYSICS

Connecting Atmospheric Science and At-
mospheric Models for Aerocaptured Mis-
sions to Titan and the Outer Planets - 5

AUTONOMOUS NAVIGATION

Micro Navigator - 23

Precision Navigation for a Mars Airplane

- 27

AUTONOMY

Precise Image-Based Motion Estimation
for Autonomous Small Body Exploration

- 30

BALLUTES

A Light-Weight Inflatable Hypersonic
Drag Device for Planetary Entry - 30

Computational Analysis of Towed Ballute
Interactions - 18

Trailing Ballute Aerocapture: Concept
and Feasibility Assessment - 12

BUOYANCY

A Light-Weight Inflatable Hypersonic
Drag Device for Planetary Entry - 30

CAMERAS

AIMS: Acousto-optic imaging spectrom-
eter for spectral mapping of solid
surfaces - 13

Optical landmark detection for spacecraft
navigation - 6

CAPTURE EFFECT

The Development and Evaluation of an
Operational Aerobraking Strategy for the
Mars 2001 Odyssey Orbiter - 17

CIRCULAR ORBITS

Aerobraking at Venus and Mars: A Com-
parison of the Magellan and Mars Global
Surveyor Aerobraking Phases - 30

COLLISIONS

Low Velocity Impact Experiments in
Microgravity - 24

COMETS

The Deep Space 4/Champollion Comet
Rendezvous and Lander Technology
Demonstration Mission - 28

COMMAND GUIDANCE

An Investigation of Terminal Guidance
and Control Techniques for a Robotic
Mars Lander - 28

COMMUNICATION SATELLITES

Daily repeat-groundtrack Mars orbits - 7

COMPUTATIONAL FLUID DYNAMICS

CFD Prediction of the BEAGLE 2 Mars
Probe Aerodynamic Database - 19

Plume Modeling and Application to Mars
2001 Odyssey Aerobraking - 16

COMPUTERIZED SIMULATION

Autonomous Aerobraking at Mars - 16

Computational Analysis of Towed Ballute
Interactions - 18

Development of a Monte Carlo Mars-
gram model for 2001 Mars Odyssey
aerobraking simulations - 15

Mars Exploration Rover Terminal De-
scent Mission Modeling and Simulation

- 9

Multibody Parachute Flight Simulations
for Planetary Entry Trajectories Using
'Equilibrium Points' - 16

Pitch control during autonomous aero-
braking for near-term Mars exploration

- 12

Plume Modeling and Application to Mars
2001 Odyssey Aerobraking - 16

COMPUTERS

Multibody Parachute Flight Simulations
for Planetary Entry Trajectories Using
'Equilibrium Points' - 6

CONTROL SURFACES

Control Surface and Afterbody Experi-
mental Aeroheating for a Proposed Mars
Smart Lander Aeroshell - 17

CONTROL SYSTEMS DESIGN

An Investigation of Terminal Guidance
and Control Techniques for a Robotic
Mars Lander - 28

A-2

CONTROL THEORY

An Investigation of Terminal Guidance
and Control Techniques for a Robotic
Mars Lander - 28

Angle-of-Attack-Modulated Terminal
Point Control for Neptune Aerocapture

- 9

CONVECTIVE HEAT TRANSFER

Preliminary Convective-Radiative Heat-
ing Environments for a Neptune Aero-
capture Mission - 1

CORRELATION

Thermal Analysis and Correlation of the
Mars Odyssey Spacecraft's Solar Array
During Aerobraking Operations - 17

DATA ACQUISITION

Approach navigation for the 2009 Mars
large lander - 6

DATA BASES

CFD Prediction of the BEAGLE 2 Mars
Probe Aerodynamic Database - 19

DEATH VALLEY (CA)

Remote Sensing of Evaporite Minerals in
Badwater Basin, Death Valley, at Varying
Spatial Scales and in Different Spectral
Regions - 22

DEEP SPACE

The Deep Space 4/Champollion Comet
Rendezvous and Lander Technology
Demonstration Mission - 28

DEGREES OF FREEDOM

Mars Exploration Rover Six-Degree-Of-
Freedom Entry Trajectory Analysis - 14

DESCENT

Mars Exploration Rover Terminal De-
scent Mission Modeling and Simulation

- 9

DIFFERENTIAL EQUATIONS

Multibody Parachute Flight Simulations
for Planetary Entry Trajectories Using
'Equilibrium Points' - 16

DRAG DEVICES

A Light-Weight Inflatable Hypersonic
Drag Device for Planetary Entry - 30

DRAG

Pitch control during autonomous aero-
braking for near-term Mars exploration

- 12

DUST

Low Velocity Impact Experiments in
Microgravity - 24

EARTH-MARS TRAJECTORIES

Navigation Strategy for the Mars 2001
Lander Mission - 29

The Stakes of the Aerocapture for Mis-
sions to Mars - 25

ELLIPTICAL ORBITS

Aeroheating Thermal Analysis Methods
for Aerobraking Mars Missions - 4

ENERGY DISSIPATION

Low Velocity Impact Experiments in
Microgravity - 24

ENVIRONMENT PROTECTION

International Agreement on Planetary
Protection - 24

ENVIRONMENTAL MONITORING

Atmospheric Models for Aerocapture - 4

ESTIMATING

Precise Image-Based Motion Estimation
for Autonomous Small Body Exploration
- 30

EUROPEAN SPACE PROGRAMS

Beagle 2: The Next Exobiology Mission
to Mars - 21

EVALUATION

Mars Smart Lander Parachute Simula-
tion Model - 19

EXOBIOLOGY

Beagle 2: The Next Exobiology Mission
to Mars - 21

FEASIBILITY ANALYSIS

Trailing Ballute Aerocapture: Concept
and Feasibility Assessment - 12

FLIGHT CONTROL

Computational Analysis of Towed Ballute
Interactions - 18

FLIGHT OPERATIONS

Autonomous Aerobraking at Mars

16

The Development and Evaluation of an
Operational Aerobraking Strategy for the
Mars 2001 Odyssey Orbiter - 17

FLIGHT SIMULATION

Mars Smart Lander Parachute Simula-
tion Model - 19

Multibody Parachute Flight Simulations
for Planetary Entry Trajectories Using
'Equilibrium Points' - 16

FLOATS

A Light-Weight Inflatable Hypersonic
Drag Device for Planetary Entry - 30

FLOW CHARACTERISTICS

Wake Closure Characteristics and After-
body Heating on a Mars Sample Return
Orbiter- 15

FLOW DISTRIBUTION

Plume Modeling and Application to Mars
2001 Odyssey Aerobraking - 16

GALILEAN SATELLITES

Science and Engineering Potential of an
Icy Moon Lander - 13

GALILEO SPACECRAFT

Pioneer Venus and Galileo entry probe
heritage - 8

GAS GIANT PLANETS

Radioisotope Electric Propulsion for Fast
Outer Planetary Orbiters - 20

GRAND TOURS

Radioisotope Electric Propulsion for Fast
Outer Planetary Orbiters - 20

GRAVITATIONAL EFFECTS

Low Velocity Impact Experiments in
Microgravity - 24

GROUND TRACKS

Daily repeat-groundtrack Mars orbits - 7

GUIDANCE (MOTION)

Atmospheric Models for Aeroentry and
Aeroassist - 1

HEAT FLUX

Autonomous Aerobraking at Mars

16

HEAT SHIELDING

Thermal protection system technology
and facility needs for demanding future
planetary missions - 8

HEAT TRANSFER COEFFICIENTS

Thermal Analysis and Correlation of the
Mars Odyssey Spacecraft's Solar Array
During Aerobraking Operations - 17

HIGH RESOLUTION

Mars reconnaissance orbiter design ap-
proach for high-resolution surface
imaging - 12

HUYGENS PROBE

Atmospheric Models for Aerocapture - 4

HYPERBOLIC TRAJECTORIES

Aerocapture Guidance Methods for High
Energy Trajectories - 11

HYPERSONIC FLOW

Experimental Hypersonic Aerodynamic
Characteristics of the 2001 Mars Sur-
veyor Precision Lander with Flap - 18

HYPERSONIC WAKES

Wake Closure Characteristics and After-
body Heating on a Mars Sample Return
Orbiter- 15

IMAGE ANALYSIS

Precise Image-Based Motion Estimation
for Autonomous Small Body Exploration

- 30

IMAGING TECHNIQUES

AIMS: Acousto-optic imaging spectrom-
eter for spectral mapping of solid
surfaces - 13

Mars reconnaissance orbiter design ap-
proach for high-resolution surface
imaging - 12

IMPACT

Low Velocity Impact Experiments in
Microgravity - 24

IMPINGEMENT

Wake Closure Characteristics and After-
body Heating on a Mars Sample Return
Orbiter- 15

IN SITU MEASUREMENT

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system

- 10

A-3

INFLATABLE STRUCTURES

A Light-Weight Inflatable Hypersonic
Drag Device for Planetary Entry - 30

INTERNATIONAL COOPERATION

International Agreement on Planetary
Protection - 24

INTERNATIONAL LAW

International Agreement on Planetary
Protection - 24

INTERNATIONAL SPACE STATION

Aeroassist Technology Planning for
Exploration - 4

INTERORBITAL TRAJECTORIES

Optimization of Low Thrust Trajectories
With Terminal Aerocapture - 11

INTERPLANETARY NAVIGATION

Navigation Strategy for the Mars 2001
Lander Mission - 29

INTERPLANETARY SPACECRAFT

Blended control, predictor-corrector guid-
ance algorithm: An enabling technology
for Mars aerocapture - 10

Wake Closure Characteristics and After-
body Heating on a Mars Sample Return
Orbiter- 15

INTERPLANETARY TRAJECTORIES

Optimization of Low Thrust Trajectories
With Terminal Aerocapture - 11

INTERPLANETARY TRANSFER ORBITS

Earth Return Aerocapture for the
TransHab/Ellipsled Vehicle - 28

INVISCID FLOW

Experimental Hypersonic Aerodynamic
Characteristics of the 2001 Mars Sur-
veyor Precision Lander with Flap - 18

ION PROPULSION

SEP Mission to Titan NEXT Aerocapture
In-Space Propulsion (Quicktime Movie)
- 10

KALMAN FILTERS

Entry trajectory and atmosphere recon-
struction methodologies for the mars ex-
ploration rover mission - 7

LANDING MODULES

Mars Exploration Rovers Landing Dis-
persion Analysis - 3

LASER ALTIMETERS

MOLA-Based Landing

Characterization - 27

Site

LANDING SITES

MOLA-Based
Characterization

Landing
27

Site

Navigation Strategy for the Mars 2001
Lander Mission - 29

LANDING

The Deep Space 4/Champollion Comet
Rendezvous and Lander Technology
Demonstration Mission - 28

LAND

Mars Exploration Rovers Landing Dis-
persion Analysis - 3

LOADS (FORCES)

Mars Smart Lander Parachute Simula-
tion Model - 19

LOW SPEED

Low Velocity Impact Experiments in
Microgravity - 24

MAGELLAN SPACECRAFT (NASA)

Aerobraking at Venus and Mars: A Com-
parison of the Magellan and Mars Global
Surveyor Aerobraking Phases - 30

MANNED MARS MISSIONS

Aerothermal Instrumentation Loads To
Implement Aeroassist Technology in Fu-
ture Robotic and Human Missions to
MARS and Other Locations Within the
Solar System - 20

MAPPING

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

MARS ATMOSPHERE

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

MARS EXPLORATION

After the Mars Polar Lander: Where to
Next? - 26

Mars Exploration Rover Six-Degree-Of-
Freedom Entry Trajectory Analysis - 1 4

Mars Exploration Rover Terminal De-
scent Mission Modeling and Simulation

- 9

Mars Exploration Rovers Landing Dis-
persion Analysis - 3

Precision Navigation for a Mars Airplane

- 27

The Martian Oasis Detector - 26

MARS GLOBAL SURVEYOR

Aerobraking at Venus and Mars: A Com-
parison of the Magellan and Mars Global
Surveyor Aerobraking Phases- 30

MOLA-Based
Characterization

Landing
27

Site

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

MARS LANDING

Control Surface and Afterbody Experi-
mental Aeroheating for a Proposed Mars
Smart Lander Aeroshell - 17

Mars Exploration Rovers Entry, Descent,
and Landing Trajectory Analysis - 3

Mars Smart Lander Parachute Simula-
tion Model - 19

Navigation Strategy for the Mars 2001
Lander Mission - 29

Precision Terminal Guidance for a Mars
Lander - 25

MARS MISSIONS

Aeroheating Thermal Analysis Methods
for Aerobraking Mars Missions - 4

After the Mars Polar Lander: Where to
Next? - 26

Autonomous Aerobraking at Mars - 16

Beagle 2: The Next Exobiology Mission
to Mars - 21

Lunar and Planetary Science XXXV: Mis-
sions and Instruments: Hopes and Hope
Fulfilled - 5

Precision Navigation for a Mars Airplane

- 27

The Martian Oasis Detector - 26

The Stakes of the Aerocapture for Mis-
sions to Mars - 25

MARS (PLANET)

After the Mars Polar Lander: Where to
Next? - 26

Precision Navigation for a Mars Airplane

- 27

The Martian Oasis Detector - 26

MARS POLAR LANDER

After the Mars Polar Lander: Where to
Next? - 27

MARS ROVING VEHICLES

Mars Exploration Rover Six-Degree-Of-
Freedom Entry Trajectory Analysis - 1 4

Mars Exploration Rover Terminal De-
scent Mission Modeling and Simulation

- 9

MARS SAMPLE RETURN MISSIONS

Mars Sample Return without Landing on
the Surface - 25

Wake Closure Characteristics and After-
body Heating on a Mars Sample Return
Orbiter- 15

MARS SURFACE

MOLA-Based
Characterization

The Martian Oasis Detector

Landing
27

Site

26

MARS SURVEYOR 2001 MISSION

Navigation Strategy for the Mars 2001
Lander Mission - 29

MATHEMATICAL MODELS

Mars Exploration Rover Terminal De-
scent Mission Modeling and Simulation
- 9

Plume Modeling and Application to Mars
2001 Odyssey Aerobraking - 16

Uncertainty Optimization Applied to the
Monte Carlo Analysis of Planetary Entry
Trajectories - 23

MEASURING INSTRUMENTS

Science and Engineering Potential of an
Icy Moon Lander - 13

MECHANICAL OSCILLATORS

Ultra-stable oscillators for planetary entry
probes - 8

A-4

METEORITES

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

MICROGRAVITY

A Light-Weight Inflatable Hypersonic
Drag Device for Planetary Entry - 30

Low Velocity Impact Experiments in
Microgravity - 24

MINERAL DEPOSITS

Remote Sensing of Evaporite Minerals in
Badwater Basin, Death Valley, at Varying
Spatial Scales and in Different Spectral
Regions - 22

MINIATURIZATION

Micro Navigator -

23

MISSION PLANNING

After the Mars Polar Lander: Where to
Next? - 26

The Deep Space 4/Champollion Comet
Rendezvous and Lander Technology
Demonstration Mission - 28

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

MONTE CARLO METHOD

Development of a Monte Carlo Mars-
gram model for 2001 Mars Odyssey
aerobraking simulations - 15

Engineering-Level Model Atmospheres
for Titan & Neptune - 13

Uncertainty Optimization Applied to the
Monte Carlo Analysis of Planetary Entry
Trajectories - 23

MULTISENSOR FUSION

Micro Navigator - 23

NASA SPACE PROGRAMS

Aeroassist Technology Planning for
Exploration - 4

Mars Exploration Rover Six-Degree-Of-
Freedom Entry Trajectory Analysis - 14

NASA Development of Aerocapture
Technologies - 14

NAVIGATION

Approach navigation for the 2009 Mars
large lander - 6

Approaches to autonomous aerobraking
at Mars - 1 5

Optical landmark detection for spacecraft
navigation - 6

The Development and Evaluation of an
Operational Aerobraking Strategy for the
Mars 2001 Odyssey Orbiter - 17

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

NEPTUNE ATMOSPHERE

Preliminary Convective-Radiative Heat-
ing Environments for a Neptune Aero-
capture Mission - 1

NEPTUNE (PLANET)

Aerocapture Performance Analysis for a
Neptune-Triton Exploration Mission - 2

Angle-of-Attack-Modulated Terminal
Point Control for Neptune Aerocapture

- 9

Engineering-Level Model Atmospheres
for Titan & Neptune - 13

Neptune Aerocapture Systems Analysis

- 3

Structural Design for a Neptune Aerocap-
ture Mission - 2

NUCLEAR ELECTRIC PROPULSION

Radioisotope Electric Propulsion for Fast
Outer Planetary Orbiters - 20

NUMERICAL ANALYSIS

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

OASES

The Martian Oasis Detector - 26

OPTICAL TRACKING

Precise Image-Based Motion Estimation
for Autonomous Small Body Exploration

- 30

OPTIMAL CONTROL

Optimization of Low Thrust Trajectories
With Terminal Aerocapture - 11

ORBIT CALCULATION

Navigation Strategy for the Mars 2001
Lander Mission - 29

ORBIT INSERTION

Aerocapture Technology Development
Needs for Outer Planet Exploration - 20

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

ORBITAL MECHANICS

Navigation Strategy for the Mars 2001
Lander Mission - 29

ORBITS

Daily repeat-groundtrack Mars orbits - 7

Mars reconnaissance orbiter design ap-
proach for high-resolution surface
imaging - 12

OUTER PLANETS EXPLORERS

Aerocapture Technology Development
Needs for Outer Planet Exploration - 20

PARACHUTES

Entry trajectory and atmosphere recon-
struction methodologies for the mars ex-
ploration rover mission - 7

Mars Smart Lander Parachute Simula-
tion Model - 19

Multibody Parachute Flight Simulations
for Planetary Entry Trajectories Using
'Equilibrium Points' - 6

PARTICULATE SAMPLING

Mars Sample Return without Landing on
the Surface - 25

PERIODIC VARIATIONS

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

PLANETARY ATMOSPHERES

A Light-Weight Inflatable Hypersonic
Drag Device for Planetary Entry - 30

Atmospheric Models for Aerocapture
Systems Studies - 1

Connecting Atmospheric Science and At-
mospheric Models for Aerocaptured Mis-
sions to Titan and the Outer Planets - 5

PLANETARY ENVIRONMENTS

International Agreement on Planetary
Protection - 24

PLANETARY LANDING

AIMS: Acousto-optic imaging spectrom-
eter for spectral mapping of solid
surfaces - 13

Approach navigation for the 2009 Mars
large lander - 6

Daily repeat-groundtrack Mars orbits - 7

Entry descent, and landing scenario for
the Mars exploration Rover mission - 7

Entry trajectory and atmosphere recon-
struction methodologies for the mars ex-
ploration rover mission - 7

Europa Lander - 13

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system
- 10

Science and Engineering Potential of an
Icy Moon Lander - 13

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Thermal protection system technology
and facility needs for demanding future
planetary missions - 8

PLANETOLOGY

Lunar and Planetary Science XXXV: Mis-
sions and Instruments: Hopes and Hope
Fulfilled - 5

PLANETS

Approaches to autonomous aerobraking
at Mars - 1 5

Blended control, predictor-corrector guid-
ance algorithm: An enabling technology
for Mars aerocapture - 10

Daily repeat-groundtrack Mars orbits - 7

Development of a Monte Carlo Mars-
gram model for 2001 Mars Odyssey
aerobraking simulations - 15

Entry descent, and landing scenario for
the Mars exploration Rover mission - 7

Pioneer Venus and Galileo entry probe
heritage - 8

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

A-5

PLANNING

The Development and Evaluation of an
Operational Aerobraking Strategy for the
Mars 2001 Odyssey Orbiter - 17

POLAR REGIONS

After the Mars Polar Lander: Where to
Next? - 26

POLICIES

International Agreement on Planetary
Protection - 24

POSITION (LOCATION)

Mars Exploration Rovers Entry, Descent,
and Landing Trajectory Analysis - 3

POWERED LIFT AIRCRAFT

Exploration of Titan Using Vertical Lift
Aerial Vehicles - 23

PREDICTIONS

CFD Prediction of the BEAGLE 2 Mars
Probe Aerodynamic Database - 19

PREDICTOR-CORRECTOR METHODS

Aerocapture Guidance Methods for High
Energy Trajectories - 11

Blended control, predictor-corrector guid-
ance algorithm: An enabling technology
for Mars aerocapture - 10

PRESSURE VESSELS

Pioneer Venus and Galileo entry probe
heritage - 8

PRODUCT DEVELOPMENT

Earth Return Aerocapture for the
TransHab/Ellipsled Vehicle - 28

PROPULSION

NASA Development of Aerocapture
Technologies - 14

QUARTZ

Ultra-stable oscillators for planetary entry
probes - 8

RADIATIVE HEAT TRANSFER

Preliminary Convective-Radiative Heat-
ing Environments for a Neptune Aero-
capture Mission - 1

RADIO TRACKING

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

RADIOACTIVE ISOTOPES

Radioisotope Electric Propulsion for Fast
Outer Planetary Orbiters - 20

RECONNAISSANCE AIRCRAFT

Mars reconnaissance orbiter design ap-
proach for high-resolution surface
imaging - 12

REMOTE SENSING

Remote Sensing of Evaporite Minerals in
Bad water Basin, Death Valley, at Varying
Spatial Scales and in Different Spectral
Regions - 22

RESEARCH

A Light-Weight Inflatable Hypersonic
Drag Device for Planetary Entry - 30

RESOLUTION

Optical landmark detection for spacecraft
navigation - 6

ROBOTICS

Aerothermal Instrumentation Loads To
Implement Aeroassist Technology in Fu-
ture Robotic and Human Missions to
MARS and Other Locations Within the
Solar System - 20

ROCKET EXHAUST

Plume Modeling and Application to Mars
2001 Odyssey Aerobraking - 16

ROTARY WING AIRCRAFT

Exploration of Titan Using Vertical Lift
Aerial Vehicles - 23

ROVING VEHICLES

Approach navigation for the 2009 Mars
large lander - 6

Entry descent, and landing scenario for
the Mars exploration Rover mission - 7

Mars Exploration Rovers Entry, Descent,
and Landing Trajectory Analysis - 3

Mars Exploration Rovers Landing Dis-
persion Analysis - 3

RUBIDIUM

Ultra-stable oscillators for planetary entry
probes - 8

SATELLITE SURFACES

Science and Engineering Potential of an
Icy Moon Lander - 13

SEDIMENTARY ROCKS

Remote Sensing of Evaporite Minerals in
Badwater Basin, Death Valley, at Varying
Spatial Scales and in Different Spectral
Regions - 22

SIMULATION

The Development and Evaluation of an
Operational Aerobraking Strategy for the
Mars 2001 Odyssey Orbiter - 17

SOLAR ARRAYS

Autonomous Aerobraking at Mars

16

Thermal Analysis and Correlation of the
Mars Odyssey Spacecraft's Solar Array
During Aerobraking Operations - 17

SOLAR ELECTRIC PROPULSION

SEP Mission to Titan NEXT Aerocapture
In-Space Propulsion (Quicktime Movie)
- 10

SOLAR SYSTEM

Aerothermal Instrumentation Loads To
Implement Aeroassist Technology in Fu-
ture Robotic and Human Missions to
MARS and Other Locations Within the
Solar System - 20

Atmospheric Models for Aeroentry and
Aeroassist - 1

Europa Lander - 13

SPACE EXPLORATION

Aeroassist Technology Planning for
Exploration - 4

Aerocapture Performance Analysis for a
Neptune-Triton Exploration Mission - 2

Exploration of Titan Using Vertical Lift
Aerial Vehicles - 23

SPACE FLIGHT

Approaches to autonomous aerobraking
at Mars - 1 5

Development of a Monte Carlo Mars-
gram model for 2001 Mars Odyssey
aerobraking simulations - 15

SPACE MISSIONS

Aerocapture Performance Analysis for a
Neptune-Triton Exploration Mission - 2

Mars Exploration Rover Terminal De-
scent Mission Modeling and Simulation
- 9

Structural Design for a Neptune Aerocap-
ture Mission - 2

Study of Orbital Transfers with Aeroas-
sisted Maneuvers - 21

The Deep Space 4/Champollion Comet
Rendezvous and Lander Technology
Demonstration Mission - 28

SPACE NAVIGATION

Micro Navigator - 23

SPACE PROBES

Beagle 2: The Next Exobiology Mission
to Mars - 21

Pioneer Venus and Galileo entry probe
heritage - 8

Science and Engineering Potential of an
Icy Moon Lander - 13

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Thermal protection system technology
and facility needs for demanding future
planetary missions - 8

SPACECRAFT CONTROL

Computational Analysis of Towed Ballute
Interactions - 18

SPACECRAFT DESIGN

A Light-Weight Inflatable Hypersonic
Drag Device for Planetary Entry - 30

Aerocapture Technology Development
Needs for Outer Planet Exploration - 20

Earth Return Aerocapture for the
TransHab/Ellipsled Vehicle - 28

Structural Design for a Neptune Aerocap-
ture Mission - 2

SPACECRAFT GUIDANCE

Aerocapture Guidance Algorithm Com-
parison Campaign - 18

Micro Navigator - 23

Precision Terminal Guidance for a Mars
Lander - 25

SPACECRAFT INSTRUMENTS

Beagle 2: The Next Exobiology Mission
to Mars - 21

A-6

Lunar and Planetary Science XXXV: Mis-
sions and Instruments: Hopes and Hope
Fulfilled - 5

The Martian Oasis Detector - 26

SPACECRAFT MANEUVERS

Aerocapture Technology Development
Needs for Outer Planet Exploration - 20

Autonomous Aerobraking at Mars - 16

SPACECRAFT MOTION

Precise Image-Based Motion Estimation
for Autonomous Small Body Exploration

- 30

SPACECRAFT PERFORMANCE

Aerocapture Performance Analysis for a
Neptune-Triton Exploration Mission - 2

SPACECRAFT PROPULSION

Aerocapture Technology Project
Overview - 11

Optical landmark detection for spacecraft
navigation - 6

SPACECRAFT

Approaches to autonomous aerobraking
at Mars - 1 5

Europa Lander - 13

Multibody Parachute Flight Simulations
for Planetary Entry Trajectories Using
'Equilibrium Points' - 6

Pitch control during autonomous aero-
braking for near-term Mars exploration

- 12

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system

- 10

SPATIAL RESOLUTION

Remote Sensing of Evaporite Minerals in
Badwater Basin, Death Valley, at Varying
Spatial Scales and in Different Spectral
Regions - 22

SPECTRAL BANDS

Remote Sensing of Evaporite Minerals in
Badwater Basin, Death Valley, at Varying
Spatial Scales and in Different Spectral
Regions - 22

STATISTICAL ANALYSIS

Mars Exploration Rovers Landing Dis-
persion Analysis - 3

STIFFNESS

Multibody Parachute Flight Simulations
for Planetary Entry Trajectories Using
'Equilibrium Points' - 6

STRUCTURAL BASINS

Remote Sensing of Evaporite Minerals in
Badwater Basin, Death Valley, at Varying
Spatial Scales and in Different Spectral
Regions - 22

STRUCTURAL DESIGN

Structural Design for a Neptune Aerocap-
ture Mission - 2

SURFACE PROPERTIES

Science and Engineering Potential of an
Icy Moon Lander - 13

SYSTEMS ANALYSIS

Aerocapture Performance Analysis for a
Neptune-Triton Exploration Mission - 2

Aerocapture Technology Project
Overview - 11

NASA Development of Aerocapture
Technologies - 14

Neptune Aerocapture Systems Analysis

- 3

TECHNOLOGICAL FORECASTING

Aeroassist Technology Planning for
Exploration - 4

TECHNOLOGY UTILIZATION

Aerothermal Instrumentation Loads To
Implement Aeroassist Technology in Fu-
ture Robotic and Human Missions to
MARS and Other Locations Within the
Solar System - 20

Neptune Aerocapture Systems Analysis

- 3

TEMPERATURE MEASURING INSTRU-
MENTS

Aerothermal Instrumentation Loads To
Implement Aeroassist Technology in Fu-
ture Robotic and Human Missions to
MARS and Other Locations Within the
Solar System - 20

TEMPERATURE PROFILES

Autonomous Aerobraking at Mars

16

TERMINAL GUIDANCE

An Investigation of Terminal Guidance
and Control Techniques for a Robotic
Mars Lander - 28

Angle-of-Attack-Modulated Terminal
Point Control for Neptune Aerocapture
- 9

Precision Terminal Guidance for a Mars
Lander - 25

THERMAL ANALYSIS

Aeroheating Thermal Analysis Methods
for Aerobraking Mars Missions - 4

Thermal Analysis and Correlation of the
Mars Odyssey Spacecraft's Solar Array
During Aerobraking Operations - 17

THRUST VECTOR CONTROL

An Investigation of Terminal Guidance
and Control Techniques for a Robotic
Mars Lander - 28

TITAN

Engineering-Level Model Atmospheres
for Titan & Neptune - 13

TOPOGRAPHY

Entry descent, and landing scenario for
the Mars exploration Rover mission - 7

TOWED BODIES

Computational Analysis of Towed Ballute
Interactions - 18

TRACKING (POSITION)

Optical landmark detection for spacecraft
navigation - 6

TRAJECTORIES

Aerocapture Guidance Methods for High
Energy Trajectories - 11

Approach navigation for the 2009 Mars
large lander - 6

Entry trajectory and atmosphere recon-
struction methodologies for the mars ex-
ploration rover mission - 7

Multibody Parachute Flight Simulations
for Planetary Entry Trajectories Using
'Equilibrium Points' - 6

The Development and Evaluation of an
Operational Aerobraking Strategy for the
Mars 2001 Odyssey Orbiter - 17

TRAJECTORY ANALYSIS

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

Mars Exploration Rover Six-Degree-Of-
Freedom Entry Trajectory Analysis - 1 4

Mars Exploration Rover Terminal De-
scent Mission Modeling and Simulation
- 9

Mars Exploration Rovers Entry, Descent,
and Landing Trajectory Analysis - 3

Mars Smart Lander Parachute Simula-
tion Model - 19

Uncertainty Optimization Applied to the
Monte Carlo Analysis of Planetary Entry
Trajectories - 23

TRAJECTORY CONTROL

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

Uncertainty Optimization Applied to the
Monte Carlo Analysis of Planetary Entry
Trajectories - 23

TRAJECTORY OPTIMIZATION

Optimization of Low Thrust Trajectories
With Terminal Aerocapture - 11

Uncertainty Optimization Applied to the
Monte Carlo Analysis of Planetary Entry
Trajectories - 23

TRAJECTORY PLANNING

Uncertainty Optimization Applied to the
Monte Carlo Analysis of Planetary Entry
Trajectories - 23

TRANSFER ORBITS

Study of Orbital Transfers with Aeroas-
sisted Maneuvers - 21

The Stakes of the Aerocapture for Mis-
sions to Mars - 25

TRANSPONDERS

Ultra-stable oscillators for planetary entry
probes - 8

TRITON

Aerocapture Performance Analysis for a
Neptune-Triton Exploration Mission - 2

VAPORIZING

Thermal protection system technology
and facility needs for demanding future
planetary missions - 8

A-7

VENUS (PLANET) WATER Control Surface and Afterbody Experi-

Pioneer Venus and Galileo entry probe The Martian Oasis Detector - 26 mental Aeroheating for a Proposed Mars

heritage- 8 Smart Lander Aeroshell - 17

VERTICAL TAKEOFF AIRCRAFT WIND TUNNEL TESTS Wake aosure characteristics and After-
Exploration of Titan Using Vertical Lift CFD Prediction of the BEAGLE 2 Mars body Heating on a Mars Sample Return
Aerial Vehicles - 23 Probe Aerodynamic Database - 19 Orbiter- 15

A-8

Corporate Sources

Air Force Inst, of Tech.

Aerocapture Guidance Methods for High
Energy Trajectories - 11

Arizona Univ.

The Martian Oasis Detector -

26

Ball Aerospace and Technologies Corp.

Trailing Ballute Aerocapture: Concept
and Feasibility Assessment - 12

California Univ.

After the Mars Polar Lander: Where to
Next? - 26

Centre National d'Etudes Spatiales

The Stakes of the Aerocapture for Mis-
sions to Mars - 25

CFD Research Corp.

CFD Prediction of the BEAGLE 2 Mars
Probe Aerodynamic Database - 19

Colorado Univ.

Low Velocity Impact Experiments in
Microgravity - 24

Computer Sciences Corp.

Atmospheric Models for Aerocapture
Systems Studies - 1

Connecting Atmospheric Science and At-
mospheric Models for Aerocaptured Mis-
sions to Titan and the Outer Planets - 5

Georgia Inst, of Tech.

Uncertainty Optimization Applied to the
Monte Carlo Analysis of Planetary Entry
Trajectories - 23

Instituto Nacional de Pesquisas Espacias

Study of Orbital Transfers with Aeroas-
sisted Maneuvers - 21

Jet Propulsion Lab., California Inst, of
Tech.

A Light-Weight Inflatable Hypersonic
Drag Device for Planetary Entry - 30

Aerobraking at Venus and Mars: A Com-
parison of the Magellan and Mars Global
Surveyor Aerobraking Phases - 30

Mars Exploration Rovers Landing Dis-
persion Analysis - 3

Mars Sample Return without Landing on
the Surface - 25

Micro Navigator - 23

MOLA-Based
Characterization

Landing
27

Site

Navigation Strategy for the Mars 2001
Lander Mission - 29

Precise Image-Based Motion Estimation
for Autonomous Small Body Exploration
- 30

The Deep Space 4/Champollion Comet
Rendezvous and Lander Technology
Demonstration Mission - 28

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

Lunar and Planetary Inst.

Lunar and Planetary Science XXXV: Mis-
sions and Instruments: Hopes and Hope
Fulfilled - 5

Morgan Research Corp.

Atmospheric Models for Aerocapture - 4

Atmospheric Models for Aeroentry and
Aeroassist - 1

NASA Ames Research Center

Aerocapture Technology Development
Needs for Outer Planet Exploration - 20

Exploration of Titan Using Vertical Lift
Aerial Vehicles - 23

International Agreement on Planetary
Protection - 24

NASA Glenn Research Center

Radioisotope Electric Propulsion for Fast
Outer Planetary Orbiters - 20

NASA Johnson Space Center

Aerocapture Guidance Algorithm Com-
parison Campaign - 18

Beagle 2: The Next Exobiology Mission
to Mars - 21

NASA Langley Research Center

Aeroassist Technology Planning for
Exploration - 4

Aerocapture Performance Analysis for a
Neptune-Triton Exploration Mission - 2

Aeroheating Thermal Analysis Methods
for Aerobraking Mars Missions - 4

Aerothermal Instrumentation Loads To
Implement Aeroassist Technology in Fu-
ture Robotic and Human Missions to
MARS and Other Locations Within the
Solar System - 20

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

Autonomous Aerobraking at Mars - 16

Computational Analysis of Towed Ballute
Interactions - 18

Control Surface and Afterbody Experi-
mental Aeroheating for a Proposed Mars
Smart Lander Aeroshell - 17

Experimental Hypersonic Aerodynamic
Characteristics of the 2001 Mars Sur-
veyor Precision Lander with Flap - 18

Mars Exploration Rover Six-Degree-Of-
Freedom Entry Trajectory Analysis - 1 4

Mars Exploration Rover Terminal De-
scent Mission Modeling and Simulation
- 9

Mars Exploration Rovers Entry, Descent,
and Landing Trajectory Analysis - 3

Mars Smart Lander Parachute Simula-
tion Model - 19

Multibody Parachute Flight Simulations
for Planetary Entry Trajectories Using
'Equilibrium Points' - 16

Neptune Aerocapture Systems Analysis

- 3

Plume Modeling and Application to Mars
2001 Odyssey Aerobraking - 16

Preliminary Convective-Radiative Heat-
ing Environments for a Neptune Aero-
capture Mission - 1

Structural Design for a Neptune Aerocap-
ture Mission - 2

The Development and Evaluation of an
Operational Aerobraking Strategy for the
Mars 2001 Odyssey Orbiter - 17

Thermal Analysis and Correlation of the
Mars Odyssey Spacecraft's Solar Array
During Aerobraking Operations - 17

Wake Closure Characteristics and After-
body Heating on a Mars Sample Return
Orbiter- 15

NASA Marshall Space Flight Center

Aerocapture Technology Project
Overview - 11

Engineering-Level Model Atmospheres
for Titan & Neptune - 13

NASA Development of Aerocapture
Technologies - 14

SEP Mission to Titan NEXT Aerocapture
In-Space Propulsion (Quicktime Movie)

- 10

Naval Postgraduate School

Optimization of Low Thrust Trajectories
With Terminal Aerocapture - 11

North Carolina State Univ.

An Investigation of Terminal Guidance
and Control Techniques for a Robotic
Mars Lander - 28

Rhode Island Univ.

Science and Engineering Potential of an
Icy Moon Lander - 13

Science Applications International Corp.

Precision Navigation for a Mars Airplane

- 27

Precision Terminal Guidance for a Mars
Lander - 25

Tennessee Univ.

Earth Return Aerocapture for
TransHab/Ellipsled Vehicle - 28

the

Remote Sensing of Evaporite Minerals in
Badwater Basin, Death Valley, at Varying
Spatial Scales and in Different Spectral
Regions - 22

B-1

Document Authors

Alter, Stephen J.

Experimental Hypersonic Aerodynamic
Characteristics of the 2001 Mars Sur-
veyor Precision Lander with Flap - 18

Wake Closure Characteristics and After-
body Heating on a Mars Sample Return
Orbiter- 15

Alwar, V.

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

Amundsen, Ruth M.

Aeroheating Thermal Analysis Methods
for Aerobraking Mars Missions - 4

Anderson, Brian P.

Computational Analysis of Towed Ballute
Interactions - 18

Asmar, S. W.

Ultra-stable oscillators for planetary entry
probes - 8

Atkinson, D. H.

Ultra-stable oscillators for planetary entry
probes - 8

Atkinson, David

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Atreya, Sushil

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Bada, J.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system

- 10

Baggett, Randy

SEP Mission to Titan NEXT Aerocapture
In-Space Propulsion (Quicktime Movie)

- 10

Baldridge, A.

Remote Sensing of Evaporite Minerals in
Bad water Basin, Death Valley, at Varying
Spatial Scales and in Different Spectral
Regions - 22

Banfield, Donald

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Beaty, D. W.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system

- 10

Beebe, Reta

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Benson, Scott

Radioisotope Electric Propulsion for Fast
Outer Planetary Orbiters - 20

Bienstock, Bernard J.

Pioneer Venus and Galileo entry probe
heritage - 8

Bird, M. K.

Ultra-stable oscillators for planetary entry
probes - 8

Birge, Brian K.

An Investigation of Terminal Guidance
and Control Techniques for a Robotic
Mars Lander - 28

Blaes, B. R.

Micro Navigator -

23

Blanchard, Robert C.

Entry trajectory and atmosphere recon-
struction methodologies for the mars ex-
ploration rover mission - 7

Blaney, Diana L.

AIMS: Acousto-optic imaging spectrom-
eter for spectral mapping of solid
surfaces - 13

Bolton, Scott

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Boynton, W. V.

After the Mars Polar Lander: Where to
Next? - 26

Braun, Robert D.

Navigation Strategy for the Mars 2001
Lander Mission - 29

Briggs, Geoffrey

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Burkhart, P. Daniel

Approach navigation for the 2009 Mars
large lander - 6

Burkhart, P. D.

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

Burnell, Simon I.

CFD Prediction of the BEAGLE 2 Mars
Probe Aerodynamic Database - 19

Chapel, J.

Mars reconnaissance orbiter design ap-
proach for high-resolution surface
imaging - 12

Charbonnier, J. M.

The Stakes of the Aerocapture for Mis-
sions to Mars - 25

Chau, S. N.

Micro Navigator -

23

Chavis, Zachary Q.

Plume Modeling and Application to Mars
2001 Odyssey Aerobraking - 16

Cheatwood, F. M.

Mars Exploration Rover Six-Degree-Of-
Freedom Entry Trajectory Analysis - 14

Cheatwood, F. McNeil

Experimental Hypersonic Aerodynamic
Characteristics of the 2001 Mars Sur-
veyor Precision Lander with Flap - 18

Cheatwood, McNeil F.

Wake Closure Characteristics and After-
body Heating on a Mars Sample Return
Orbiter- 15

Cheng, Yang

Optical landmark detection for spacecraft
navigation - 6

Cianciolo, Alicia Dwyer

Autonomous Aerobraking at Mars - 16

Cledassou, R.

The Stakes of the Aerocapture for Mis-
sions to Mars - 25

Colwell, J. E.

Low Velocity Impact Experiments in
Microgravity - 24

Conrad, P.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system
- 10

Crisp, D.

After the Mars Polar Lander: Where to
Next? - 26

Crisp, David

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Cutts, James

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Cwynar, D.

Mars reconnaissance orbiter design ap-
proach for high-resolution surface
imaging - 12

Dec, John A.

Aeroheating Thermal Analysis Methods
for Aerobraking Mars Missions - 4

Thermal Analysis and Correlation of the
Mars Odyssey Spacecraft's Solar Array
During Aerobraking Operations - 17

Dec, John

Autonomous Aerobraking at Mars - 16

DeJong, E.

After the Mars Polar Lander: Where to
Next? - 26

C-1

Delamere, A.

Mars reconnaissance orbiter design ap-
proach for high-resolution surface
imaging - 12

Demcak, S. W.

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

Desai, Prasun N.

Entry descent, and landing scenario for
the Mars exploration Rover mission - 7

Entry trajectory and atmosphere recon-
struction methodologies for the mars ex-
ploration rover mission - 7

Mars Exploration Rover Six-Degree-Of-
Freedom Entry Trajectory Analysis - 14

Mars Exploration Rovers Entry, Descent,
and Landing Trajectory Analysis - 3

Mars Exploration Rovers Landing Dis-
persion Analysis - 3

DHondt, S. L.

Science and Engineering Potential of an
Icy Moon Lander - 13

Dicarlo, Jennifer L.

Aerocapture Guidance Methods for High
Energy Trajectories - 11

Diehl, Roger

Daily repeat-groundtrack Mars orbits - 7

Drake, Michael

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Dupuis, E.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system
- 10

Duvall, Aleta L.

Atmospheric Models for Aerocapture - 4

Duvall, Aleta

Atmospheric Models for Aerocapture
Systems Studies - 1

Atmospheric Models for Aeroentry and
Aeroassist - 1

Connecting Atmospheric Science and At-
mospheric Models for Aerocaptured Mis-
sions to Titan and the Outer Planets - 5

Duxbury, T. C.

MOLA-Based
Characterization

Landing
27

Site

Duxbury, Thomas C.

Mars Exploration Rovers Landing Dis-
persion Analysis - 3

Dwyer, A. M.

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

Dwyer, Alicia M.

Development of a Monte Carlo Mars-
gram model for 2001 Mars Odyssey
aerobraking simulations - 15

Dyke, R. Eric

Structural Design for a Neptune Aerocap-
ture Mission - 2

Edquist, Karl T.

Control Surface and Afterbody Experi-
mental Aeroheating for a Proposed Mars
Smart Lander Aeroshell - 17

Ely, Todd

Daily repeat-groundtrack Mars orbits - 7

Escalera, P. E.

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

Esposito, Larry

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Esposito, P. B.

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

et al.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system
- 10

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Farmer, J.

Remote Sensing of Evaporite Minerals in
Badwater Basin, Death Valley, at Varying
Spatial Scales and in Different Spectral
Regions - 22

Galal, Kenneth

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Gasbarre, Joseph F.

Thermal Analysis and Correlation of the
Mars Odyssey Spacecraft's Solar Array
During Aerobraking Operations - 17

Gefert, Leon

Radioisotope Electric Propulsion for Fast
Outer Planetary Orbiters - 20

Gehling, R.

Mars reconnaissance orbiter design ap-
proach for high-resolution surface
imaging - 12

George, B. E.

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

George, Benjamin E.

Aeroheating Thermal Analysis Methods
for Aerobraking Mars Missions - 4

Thermal Analysis and Correlation of the
Mars Odyssey Spacecraft's Solar Array
During Aerobraking Operations - 17

Gibson, Everett K., Jr.

Beagle 2: The Next Exobiology Mission
to Mars - 21

Glenar, David A.

AIMS: Acousto-optic imaging spectrom-
eter for spectral mapping of solid
surfaces - 13

Gnoffo, Peter A.

Computational Analysis of Towed Ballute
Interactions - 18

Golombek, Matthew P.

Mars Exploration Rovers Landing Dis-
persion Analysis - 3

Graat, E. J.

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

Graves, Claude

Aerocapture Guidance Algorithm Com-
parison Campaign - 18

Aerocapture Technology Development
Needs for Outer Planet Exploration - 20

Grimes, James M.

The Deep Space 4/Champollion Comet
Rendezvous and Lander Technology
Demonstration Mission - 28

Gulick, Doug

Trailing Ballute Aerocapture: Concept
and Feasibility Assessment - 12

Habchi, Sami D.

CFD Prediction of the BEAGLE 2 Mars
Probe Aerodynamic Database - 19

Hall, Jeff

Aerocapture Technology Development
Needs for Outer Planet Exploration - 20

Hanna, J. L.

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

Approaches to autonomous aerobraking
at Mars - 1 5

Hanna, Jill L.

Autonomous Aerobraking at Mars

16

Hansen, C. J.

After the Mars Polar Lander: Where to
Next? - 26

Harri, A. M.

After the Mars Polar Lander: Where to
Next? - 26

Hillman, John J.

AIMS: Acousto-optic imaging spectrom-
eter for spectral mapping of solid
surfaces - 13

Hoffmann, C.

Earth Return Aerocapture for
TransHab/Ellipsled Vehicle - 28

the

Gershman, Robert

Europa Lander -

13

Hollis, Brian R.

Control Surface and Afterbody Experi-
mental Aeroheating for a Proposed Mars
Smart Lander Aeroshell - 17

C-2

Preliminary Convective-Radiative Heat-
ing Environments for a Neptune Aero-
capture Mission - 1

Horvath, Thomas J.

Experimental Hypersonic Aerodynamic
Characteristics of the 2001 Mars Sur-
veyor Precision Lander with Flap - 18

Wake Closure Characteristics and After-
body Heating on a Mars Sample Return
Orbiter- 15

Hrinda, Glenn A.

Structural Design for a Neptune Aerocap-
ture Mission - 2

Hubbard, William

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Hunten, Donald

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Huntsberger, T.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system
- 10

Ingersoll, Andrew

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Ivanov, A. B.

MOLA-Based

Landing

Site

Characterization -

27

Ivlev, R.

Planning for a Mars

in situ sample prepa-

ration and distribution

(SPAD)

system

- 10

James, Bonnie

Aerocapture Technology Project
Overview - 1 1

NASA Development of Aerocapture
Technologies - 14

Jits, Roman Y.

Blended control, predictor-corrector guid-
ance algorithm: An enabling technology
for Mars aerocapture - 10

Johnson, Andrew E.

Optical landmark detection for spacecraft
navigation - 6

Johnson, Andrew Edie

Precise Image-Based Motion Estimation
for Autonomous Small Body Exploration

- 30

Johnson, D. L.

Engineering-Level Model Atmospheres
for Titan & Neptune - 13

Johnson, Wyatt R.

Pitch control during autonomous aero-
braking for near-term Mars exploration

- 12

Johnston, M. D.

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

Jones, Steven M.

Mars Sample Return without Landing on
the Surface - 25

Josselyn, Scott B.

Optimization of Low Thrust Trajectories
With Terminal Aerocapture - 11

Jurewicz, A. J. G.

Mars Sample Return without Landing on
the Surface - 25

Justus, C. G.

Atmospheric Models for Aerocapture
Systems Studies - 1

Atmospheric Models for Aerocapture - 4

Atmospheric Models for Aeroentry and
Aeroassist - 1

Connecting Atmospheric Science and At-
mospheric Models for Aerocaptured Mis-
sions to Titan and the Outer Planets - 5

Engineering-Level Model Atmospheres
for Titan & Neptune - 13

Kass, David M,

Mars Exploration Rovers Landing Dis-
persion Analysis - 3

Keating, G. M.

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

Keller, H. U.

After the Mars Polar Lander: Where to
Next? - 26

Keller, Vernon W.

Atmospheric Models for Aerocapture
Systems Studies - 1

Atmospheric Models for Aerocapture - 4

Atmospheric Models for Aeroentry and
Aeroassist - 1

Connecting Atmospheric Science and At-
mospheric Models for Aerocaptured Mis-
sions to Titan and the Outer Planets - 5

Kennedy, Brian M.

Mars Exploration Rovers Landing Dis-
persion Analysis - 3

Kerridge, Stuart

Daily repeat-groundtrack Mars orbits - 7

Kia, T.

Micro Navigator - 23

Kim, S. S.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system
- 10

Klarquist, William N.

Precision Terminal Guidance for a Mars
Lander - 25

Knocke, Philip C.

Mars Exploration Rovers Entry, Descent,
and Landing Trajectory Analysis - 3

Mars Exploration Rovers Landing Dis-
persion Analysis - 3

Lam-Trong, Th.

The Stakes of the Aerocapture for Mis-
sions to Mars - 25

Laub, B.

Thermal protection system technology
and facility needs for demanding future
planetary missions - 8

Lee, B. G.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system

- 10

Lee, S. W.

Mars reconnaissance orbiter design ap-
proach for high-resolution surface
imaging - 12

Lee, Wayne J.

Entry descent, and landing scenario for
the Mars exploration Rover mission - 7

Leigh, Dennis

Beagle 2: The Next Exobiology Mission
to Mars - 21

Leshin, L. A.

After the Mars Polar Lander: Where to
Next? - 26

Lewis, Jake

Trailing Ballute Aerocapture: Concept
and Feasibility Assessment - 12

Liechty, Derek S.

Control Surface and Afterbody Experi-
mental Aeroheating for a Proposed Mars
Smart Lander Aeroshell - 17

Liever, Peter A.

CFD Prediction of the BEAGLE 2 Mars
Probe Aerodynamic Database - 19

Lindstrom, D.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system

- 10

Lingard, Steve J.

CFD Prediction of the BEAGLE 2 Mars
Probe Aerodynamic Database - 19

Lockwood, Mary Kae

Neptune Aerocapture Systems Analysis

- 3

Longuski, James M.

Pitch control during autonomous aero-
braking for near-term Mars exploration

- 12

Lorenzoni, L.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system

- 10

Lowrie, James W.

Precision Navigation for a Mars Airplane

- 27

Precision Terminal Guidance for a Mars
Lander - 25

C-3

Lyne, J. E.

Earth Return Aerocapture for
TransHab/Ellipsled Vehicle - 28

the

Lyons, Daniel T.

Aerobraking at Venus and Mars: A Com-
parison of the Magellan and Mars Global
Surveyor Aerobraking Phases - 30

Pitch control during autonomous aero-
braking for near-term Mars exploration

- 12

Trailing Ballute Aerocapture: Concept
and Feasibility Assessment - 12

Mahaffy, P.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system

- 10

Masciarelli, James P.

Aerocapture Guidance Algorithm Com-
parison Campaign - 18

Aerocapture Performance Analysis for a
Neptune-Triton Exploration Mission - 2

Mase, Robert A.

Navigation Strategy for the Mars 2001
Lander Mission - 29

Matthies, Larry H.

Optical landmark detection for spacecraft
navigation - 6

Precise Image-Based Motion Estimation
for Autonomous Small Body Exploration

- 30

May, R. D.

After the Mars Polar Lander: Where to
Next? - 26

McEwen, A.

The Martian Oasis Detector -

26

McNamara, K.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system

- 10

McRonald, Angus D.

A Light-Weight Inflatable Hypersonic
Drag Device for Planetary Entry - 30

Miller, Kevin L.

Trailing Ballute Aerocapture: Concept
and Feasibility Assessment - 12

Miller, S.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system

- 10

Millerr, J. H.

Science and Engineering Potential of an
Icy Moon Lander - 13

Moersch, J. E.

Remote Sensing of Evaporite Minerals in
Bad water Basin, Death Valley, at Varying
Spatial Scales and in Different Spectral
Regions - 22

NASA Development of Aerocapture
Technologies - 14

Morgan, G.

Beagle 2: The Next Exobiology Mission
to Mars - 21

Morse, Andy

Beagle 2: The Next Exobiology Mission
to Mars - 21

Muirhead, Brian K.

The Deep Space 4/Champollion Comet
Rendezvous and Lander Technology
Demonstration Mission - 28

Munk, Michelle M.

Aeroassist Technology Planning for
Exploration - 4

Development of a Monte Carlo Mars-
gram model for 2001 Mars Odyssey
aerobraking simulations - 15

The Development and Evaluation of an
Operational Aerobraking Strategy for the
Mars 2001 Odyssey Orbiter - 17

Munk, Michelle

Aerocapture Technology Development
Needs for Outer Planet Exploration - 20

Aerocapture Technology Project
Overview - 11

NASA Development of Aerocapture
Technologies - 14

Muth, W. D.

Earth Return Aerocapture for the
TransHab/Ellipsled Vehicle - 28

Neelon, Joseph

Daily repeat-groundtrack Mars orbits - 7

Nilsen, Erik

Europa Lander

13

Noreen, Gary

Daily repeat-groundtrack Mars orbits - 7

Oberto, Robert

Europa Lander

13

Moon, Steve

Aerocapture Technology
Overview - 1 1

Project

OConnell, Tod F.

Experimental Hypersonic Aerodynamic
Characteristics of the 2001 Mars Sur-
veyor Precision Lander with Flap - 18

Olds, John

Uncertainty Optimization Applied to the
Monte Carlo Analysis of Planetary Entry
Trajectories - 23

Olejniczak, Joseph

Preliminary Convective-Radiative Heat-
ing Environments for a Neptune Aero-
capture Mission - 1

Oleson, Steven

Radioisotope Electric Propulsion for Fast
Outer Planetary Orbiters - 20

Olson, Clark F.

Optical landmark detection for spacecraft
navigation - 6

Paige, D. A.

After the Mars Polar Lander: Where to
Next? - 26

Papanastassiou, D.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system

- 10

Parker, Timothy J.

Mars Exploration Rovers Landing Dis-
persion Analysis - 3

Parmar, Devendra S.

Aerothermal Instrumentation Loads To
Implement Aeroassist Technology in Fu-
ture Robotic and Human Missions to
MARS and Other Locations Within the
Solar System - 20

Partridge, Harry

Aerocapture Technology Development
Needs for Outer Planet Exploration - 20

Patterson, Michael

Radioisotope Electric Propulsion for Fast
Outer Planetary Orbiters - 20

Perot, Etienne

Aerocapture Guidance Algorithm Com-
parison Campaign - 18

Pillinger, Colin T.

Beagle 2: The Next Exobiology Mission
to Mars - 21

Portock, B. M.

The Strategy for the Second Phase of
Aerobraking Mars Global Surveyor - 31

Powell, Richard W.

Aeroassist Technology Planning for
Exploration - 4

Entry trajectory and atmosphere recon-
struction methodologies for the mars ex-
ploration rover mission - 7

The Development and Evaluation of an
Operational Aerobraking Strategy for the
Mars 2001 Odyssey Orbiter - 17

Powell, Richard

Aerocapture Technology Development
Needs for Outer Planet Exploration - 20

Prabhu, Dinesh

Preliminary Convective-Radiative Heat-
ing Environments for a Neptune Aero-
capture Mission - 1

Prabhu, Ramadas K.

Experimental Hypersonic Aerodynamic
Characteristics of the 2001 Mars Sur-
veyor Precision Lander with Flap - 18

Praine, Ian

Beagle 2: The Next Exobiology Mission
to Mars - 21

Queen, Eric M.

Angle-of-Attack-Modulated Terminal
Point Control for Neptune Aerocapture

- 9

Mars Exploration Rover Terminal De-
scent Mission Modeling and Simulation

- 9

C-4

Mars Smart Lander Parachute Simula-
tion Model - 19

Queen, Eric

Aerocapture Guidance Algorithm Com-
parison Campaign - 18

Raiszadeh, Behzad

Mars Exploration Rover Terminal De-
scent Mission Modeling and Simulation
- 9

Raiszadeh, Ben

Mars Smart Lander Parachute Simula-
tion Model - 19

Multibody Parachute Flight Simulations
for Planetary Entry Trajectories Using
'Equilibrium Points' - 6

Rice, J.

The Martian Oasis Detector -

26

Rousseau, Stephane

Aerocapture Guidance Algorithm Com-
parison Campaign - 18

Sabahi, Dara

The Deep Space 4/Champollion Comet
Rendezvous and Lander Technology
Demonstration Mission - 28

Schoenenberger, Mark

Mars Exploration Rover Six-Degree-Of-
Freedom Entry Trajectory Analysis - 14

Schreiber, Jeffrey

Radioisotope Electric Propulsion for Fast
Outer Planetary Orbiters - 20

Schulz, Walkiria

Study of Orbital Transfers with Aeroas-
sisted Maneuvers - 21

Shams, Qamar A.

Aerothermal Instrumentation Loads To
Implement Aeroassist Technology in Fu-
ture Robotic and Human Missions to
MARS and Other Locations Within the
Solar System - 20

Sims, Mark R.

Beagle 2: The Next Exobiology Mission
to Mars - 21

Skulsky, E. D.

Mars reconnaissance orbiter design ap-
proach for high-resolution surface
imaging - 12

Smith, John C.

Navigation Strategy for the Mars 2001
Lander Mission - 29

Smith, P. H.

After the Mars Polar Lander: Where to
Next? - 26

The Martian Oasis Detector - 26

Smythe, William D.

The Deep Space 4/Champollion Comet
Rendezvous and Lander Technology
Demonstration Mission - 28

Spencer, David A.

Navigation Strategy for the Mars 2001
Lander Mission - 29

Starr, Brett R.

Aerocapture Performance Analysis for a
Neptune-Triton Exploration Mission - 2

Stein, Jim

Trailing Ballute Aerocapture: Concept
and Feasibility Assessment - 12

Stewart, Jenny

Beagle 2: The Next Exobiology Mission
to Mars - 21

Sture, S.

Low Velocity Impact Experiments in
Microgravity - 24

Sutton, Kenneth

Preliminary Convective-Radiative Heat-
ing Environments for a Neptune Aero-
capture Mission - 1

Takashima, Naruhisa

Preliminary Convective-Radiative Heat-
ing Environments for a Neptune Aero-
capture Mission - 1

Tan-Wang, Grace H.

The Deep Space 4/Champollion Comet
Rendezvous and Lander Technology
Demonstration Mission - 28

Tartabini, Paul V.

Development of a Monte Carlo Mars-
gram model for 2001 Mars Odyssey
aerobraking simulations - 15

The Development and Evaluation of an
Operational Aerobraking Strategy for the
Mars 2001 Odyssey Orbiter - 17

Tolson, R. H.

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

Approaches to autonomous aerobraking
at Mars - 15

Tolson, Robert H.

Development of a Monte Carlo Mars-
gram model for 2001 Mars Odyssey
aerobraking simulations - 15

Tolson, Robert

Autonomous Aerobraking at Mars - 16

tomasko, M. G.

The Martian Oasis Detector - 26

Trochman, Bill

Trailing Ballute Aerocapture: Concept
and Feasibility Assessment - 12

Turner, Andrew E.

Daily repeat-groundtrack Mars orbits - 7

Venkatapathy, E.

Thermal protection system technology
and facility needs for demanding future
planetary missions - 8

Wahl, Beth E.

Precision Terminal Guidance for a Mars
Lander - 25

Walberg, Gerald D.

Blended control, predictor-corrector guid-
ance algorithm: An enabling technology
for Mars aerocapture - 10

Walberg, Gerald

An Investigation of Terminal Guidance
and Control Techniques for a Robotic
Mars Lander - 28

Wawrzyniak, Geoffrey G.

Mars Exploration Rovers Landing Dis-
persion Analysis - 3

Way, David

Uncertainty Optimization Applied to the
Monte Carlo Analysis of Planetary Entry
Trajectories - 23

Weissman, Paul R.

The Deep Space 4/Champollion Comet
Rendezvous and Lander Technology
Demonstration Mission - 28

Wercinski, Paul

Aerocapture Technology Development
Needs for Outer Planet Exploration - 20

Werner, M. R.

Application of Accelerometer Data to
Mars Odyssey Aerobraking and Atmo-
spheric Modeling - 19

Westhelle, Carlos H.

Aerocapture Performance Analysis for a
Neptune-Triton Exploration Mission - 2

Wilmoth, Richard G.

Plume Modeling and Application to Mars
2001 Odyssey Aerobraking - 16

Trailing Ballute Aerocapture: Concept
and Feasibility Assessment - 12

Wake Closure Characteristics and After-
body Heating on a Mars Sample Return
Orbiter- 15

Wood, G. E.

Ultra-stable oscillators for planetary entry
probes - 8

Wright, Ian P.

Beagle 2: The Next Exobiology Mission
to Mars - 21

Wright, Michael J.

Preliminary Convective-Radiative Heat-
ing Environments for a Neptune Aero-
capture Mission - 1

Yen, A. S.

Mars Sample Return without Landing on
the Surface - 25

Young, L. A.

Exploration of Titan Using Vertical Lift
Aerial Vehicles - 23

Young, Richard E.

Summary of the Boulder Entry Probe
Workshop April 21-22, 2003, Boulder,
Colorado, USA - 9

Zimmerman, W.

Planning for a Mars in situ sample prepa-
ration and distribution (SPAD) system
- 10

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